1. An aerial missile including, in combination, an airframe comprising a plurality of connected body assemblies having duct sections defining a duct extending throughout the length of the body and having a wall, said wall providing an inlet, a diffuser communicating with the inlet, a combustion chamber communicating with the diffuser and an exit nozzle communicating with the combustion chamber, a cowl lip connecting the forwardmost body assembly and the wall of the forwardmost duct section at their corresponding forward ends, wings on the airframe and mounted for rocking movement about their root axes, longitudinally spaced partitions surrounding the duct wall and defining a compartment, an explosive charge in the compartment and surrounding the duct, a fuze system for the charge, and means mounted on one of the partitions and operable for imparting rocking movement to the wings.

Current U.S. Class:	244/3.21 ; 60/768
Current International Class:	F02C 7/22 (20060101); F02K 7/00 (20060101); F23R 3/00 (20060101); F02K 9/95 (20060101); F02K 9/00 (20060101); F02K 7/18 (20060101); F23R 3/60 (20060101); F41F 3/00 (20060101); F41F 3/04 (20060101); F42B 10/00 (20060101); F41G 7/20 (20060101); F42B 10/64 (20060101); F41G 7/00 (20060101); H01Q 1/28 (20060101); H01Q 1/27 (20060101); F42B 015/02 ()
Field of Search:	244/14,3.21 114/20.1 343/7,7.3 102/50,49 60/270

Sciascia; R. S. Lacey; J. S. Finch; W. G.

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Claims

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What is claimed is:

1. An aerial missile including, in combination, an airframe comprising a plurality of connected body assemblies having duct sections defining a duct extending throughout the length of the body and having a wall, said wall providing an inlet, a diffuser communicating with the inlet, a combustion chamber communicating with the diffuser and an exit nozzle communicating with the combustion chamber, a cowl lip connecting the forwardmost body assembly and the wall of the forwardmost duct section at their corresponding forward ends, wings on the airframe and mounted for rocking movement about their root axes, longitudinally spaced partitions surrounding the duct wall and defining a compartment, an explosive charge in the compartment and surrounding the duct, a fuze system for the charge, and means mounted on one of the partitions and operable for imparting rocking movement to the wings.

2. An aerial missile as recited in claim 1, wherein said means includes a plurality of hydraulically operated wing actuators, one of said actuators being provided for each said wing, a source of hydraulic fluid under pressure connected to said actuators, means for controlling flow of fluid to said actuators in response to an electrical signal, and links mechanically connecting the actuators to the wings.

3. In combination with an aerial missile having an airframe including a center body assembly, wings mounted for rocking movement on said body assembly, and means for boosting the missile to supersonic speed; hydraulically operated means within the missile and operable for imparting movement to the wings, said hydraulically operated means including a source of hydraulic fluid, a pump, a plurality of wing actuators, a turbine for operating the pump, means for supplying air under pressure to the turbine during flight of the missile, linkages connecting the actuators with the wings, valve means for controlling fluid flow to the actuators, and electrically operated means for controlling the valve means.

4. In combination with an aerial missile having an airframe comprising connected body assemblies, wings mounted for rocking movement on one of the body assemblies, guidance apparatus in the missile and producing electrical voltages representing missile position errors, and booster means for accelerating the missile to supersonic speed; hydraulically operated means within the missile and operable for imparting movement to the wings, said means including a source of hydraulic fluid under pressure within the missile, a pump, a plurality of wing actuators, a turbine for operating the pump, means for supplying air under pressure to the turbine during flight of the missile, linkages connecting the actuators with the wings, valve means for controlling fluid flow to the actuators, and electrically operated means connected with the guidance apparatus and responsive to said electrical voltages for controlling said valve means, whereby said actuators will be operated by said fluid flow for shifting the wings to correct for said position errors.

5. The combination recited in claim 3, including means for locking the wings during operation of the booster means, said wing locking means being releasable upon termination of operation of the booster means.

6. The combination recited in claim 3, including additionally means for locking the wings during operation of the booster means, said wing locking means including a cylinder mounted in the missile adjacent each wing, said wing having a recess, a locking pin having a piston movable in the cylinder, a spring normally urging one end of the pin into the recess for locking the wing, a solenoid operated valve, fluid connections between the solenoid operated valve, the cylinder and the hydraulic means, a source of electric power for the solenoid valve, and means for connecting said electric power source to the solenoid of the solenoid operated valve for opening said valve upon separation of the booster means from the missile whereby fluid will flow to the cylinder for shifting the piston against the compression of the spring and withdrawing the locking pin from the recess for unlocking the wing.

7. The combination recited in claim 6, including additionally means operable for releasing the wings, whereby said wings may be detached from the body assembly.

8. In an aerial missile having an airframe and a body assembly, and wings mounted for rocking movement on the body assembly; means for mounting said wings for quick detachment from the body assembly, said means including a wing mounting bulkhead in the body assembly for each wing, a wing retaining sleeve in the bulkhead, each said wing having a spar mounted in the sleeve, a bearing rotatably connecting the sleeve with the bulkhead, tongues on the sleeve, and latches on the body assembly and engageable with the tongues.

9. The combination recited in claim 4, wherein said last mentioned means includes torque motors for said valve means, each of said torque motors comprising an armature, and a pair of coils surrounding the armature, said armature being movable in one direction by energization of one of said coils and being movable in the opposite direction by energization of the other of said coils.

10. The combination recited in claim 4, wherein said missile airframe includes an inner body mounted within the forwardmost body assembly at its forward end, and said hydraulically operated means includes an accumulator mounted in said inner body.

11. An aerial missile including, in combination, an airframe including forward, center, and aft body assemblies each having a wall defining a duct section, said assemblies being connected in alignment whereby the duct sections thereof will define a continuous duct extending throughout the length of the missile, the wall of the duct being shaped to define an inlet, a diffuser, and an exit nozzle, and a combustor in the duct within the center and aft body assemblies and including an outer shroud, a central can within the outer shroud, a pilot fuel injector assembly surrounding the central can at its forward end and fuel supply means extending into the combustor and connected to the pilot fuel injector assembly.

12. An aerial missile as recited in claim 11, including additionally fuel injector means in the central can, a source of fuel in the missile and positioned about the center body assembly, means connecting the fuel source with the fuel injector means, and fuel igniter means in the combustor.

13. A hydraulic system for a missile having a movable wing and a duct, comprising, a pump in the missile, a turbine in the missile and operable by airflow passing through the duct during flight of the missile, a wing actuator in the missile and including a cylinder, a piston in the cylinder and having a piston rod, linkage means connecting the piston rod to the wing, means connecting the wing actuator to the pump, a source of hydraulic fluid for the pump, the cylinder and for said last mentioned means, said pump forcing said fluid through said last mentioned means into said cylinder for shifting said piston and rod whereby said linkage will be moved for moving the wing, and means connected between said last mentioned means and said cylinder for controlling fluid flow to the cylinder in response to an electric signal.

14. A hydraulic system as recited in claim 13, wherein said fluid flow controlling means comprises a transfer valve, and a torque motor operatively connected with the valve.

15. A hydraulic system as recited in claim 13, including additionally means for locking the wing prior to operation of the turbine and pump.

16. A hydraulic system as recited in claim 13, wherein the wing is formed with a recess at its root end, and including additionally means for locking the wing against movement prior to operation of the turbine and pump, said last mentioned means including a cylinder, a locking pin having a piston movable in the cylinder, and a spring urging the pin into the recess.

17. In a missile having an airframe including a body, and wings mounted for rocking movement on the airframe; a hydraulic system in the missile and operative for imparting rocking movements to said wings, comprising, a turbine in the airframe, a pump driven by the turbine, a sump, a pair of accumulators, hydraulic fluid in the system, wing actuators operatively connected with the wings, one said actuator being provided for each wing and each said actuator including a cylinder, a piston in the cylinder and means connecting the piston with the wing; fluid conductor means, and electrically operated valve means, said fluid conductor means being connected between the pump, accumulators, sump, wing actuators and valve means, operation of the pump causing fluid flow to the actuator cylinders under control of said valve means.

18. The structure recited in claim 17, wherein said electrically operated valve means includes a transfer valve, and a torque motor for operating the transfer valve.

19. In an aerial missile having an airframe comprising a plurality of connected body assemblies each having a duct section, said duct sections defining a continuous duct for the missile, said continuous duct including a diffuser, a combustion chamber and an exit nozzle; a combustor in the combustion chamber and including an outer shroud, a central can mounted axially within the outer shroud, a pilot fuel injector assembly mounted within the upstream end portion of the outer shroud and surrounding the upstream end of the central can and fuel supply means mounted axially of the combustion chamber and connected to the pilot fuel injector assembly.

20. A combustor for an aerial missile, comprising, an outer shroud, a central can mounted axially within the outer shroud, a pilot fuel injector assembly mounted within the outer shroud and surrounding a portion of the central can, and fuel distribution means extending into the combustor axially of the shroud assembly and connected to the pilot fuel injector assembly.

21. A combustor as recited in claim 20, including igniter means extending into the outer shroud in the vicinity of the fuel distribution means.

22. A combustor for an aerial missile, comprising, a frusto-conical outer shroud, a central can mounted in the outer shroud, a pilot fuel injector assembly mounted concentrically within the outer shroud, pilot fuel injector means connected with the pilot fuel injector, main fuel injector means, and means mounting the pilot injector means and main injector means in spaced relation to each other within the shroud.

23. A combustor for an aerial missile, comprising, a frusto-conical outer shroud having a downstream end presented away from a fluid flow, a central can within the outer shroud, a pilot fuel injector assembly mounted concentrically within the upstream end portion of the shroud and surrounding one end of the central can, fuel distribution means connected to the pilot fuel injector assembly near one end thereof, main fuel injector means within the pilot fuel injector assembly and extending downstream of the pilot fuel injector means assembly, fuel supply means for said pilot fuel injector assembly and for said main fuel injector means, and igniter means mounted in the shroud upstream of the pilot fuel injector means.

24. In combination with a aerial missile having an airframe having a diffuser, a combustor in the airframe downstream of the diffuser, and fuel injector means in the combustor; a fuel system comprising a fuel tank, first fuel pressurization means in the missile and connected with the tank for initially pressurizing fuel in the tank, a second fuel pressurization means in the missile for pressurizing fuel in the tank during flight of the missile and after decay of pressure from said first fuel pressurization means to a predetermined value, said second fuel pressurization means including a scoop mounted to receive air flow passing through the diffuser during missile flight, a pump, a turbine connected with the pump for driving said pump, means connecting the pump with the scoop, and means for connecting the pump with the fuel injector means, operation of the pump supplying fuel to the injector means during flight of the missile.

25. The combination recited in claim 24 including additionally fuel regulator means connected with the pump, said last mentioned means and the fuel injector means.

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Description

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This invention relates generally to aerial guided missiles, and more particularly to ramjet-propelled, supersonic, surface-to-air guided missiles designed to intercept and destroy aircraft carrying guided missiles.

It is an object of this invention, to provide a ramjet-propelled, supersonic, surface-to-air guided missile which is capable of intercepting and destroying high performance aircraft carrying guided missiles.

Another object of this invention is to provide a ramjet-propelled, supersonic, surface-to-air guided missile which is stabilized through the boost phase of aerial flight.

Another object of this invention is to provide an aerial guided missile which is launched directly into a programmed auxiliary beam, other than the tracking radar beam, for capture and midcourse guidance.

Another object of this invention is to provide an aerial missile which utilizes beam-rider guidance during the midcourse phase of flight which is provided by means of an auxiliary radar beam and which acts solely as a transmitter.

Even still another object to this invention is to provide an aerial guided missile in which the guidance beam elevation and azimuth are programmed so as to approach coincidence with the target-tracking radar beam prior to terminal guidance, such that the elevation program permits "up and over" trajectories that are favorable from the standpoint of fuel economy while the azimuth program avoids restrictions in firing direction arising from launching interferences.

Another object of this invention is to provide an aerial guided missile in which semiactive radar interferometer homing is utilized during the terminal phase of flight of the aerial missile to the target, with the target illumination being provided by a tracking radar and in which target discrimination is accomplished by means of the homing equipment through range gating.

Other objects and many of the attendant advantages of this invention will be readily appreciated as the same becomes better understood by reference to the following detailed description when considered in connection with the accompanying drawings, wherein:

FIG. 1 is a side elevation view of the right hand side of a missile according to the invention and showing the location of the various systems;

FIG. 2 is a front elevation view of the missile showing the arrangement of the wings;

FIG. 3 is a rear elevation view of the missile;

FIG. 4 is an elevation view of the right hand side of the missile shown with the outer skin removed from certain system compartments thereof;

FIG. 5 is a cross sectional view taken on line 5--5 of FIG. 4, and shows the location of the three intelligence systems thereof packages;

FIG. 6 is a sectional view taken on line 6--6 of FIG. 4 and shows the method of mounting the two warhead halves;

FIG. 7 is a fragmentary vertical sectional view taken through the forward end of the missile and the inner body thereof;

FIG. 8 is a fragmentary elevation view of the fuel system portion of the missile and showing the packaging arrangement of the various fuel system components;

FIGS. 9A and 9B considered together, comprise a diagrammatic illustration of the fuel system for the missile;

FIG. 10 is a diagrammatic sectional view of the main fuel regulator;

FIG. 11 is a diagrammatic sectional view of the pilot regulator of the fuel system;

FIG. 12 is a diagrammatic view of a section the turbine fuel pump of the fuel system;

FIG. 13 is a diagrammatic view in section of a fuel shutoff valve of the fuel system;

FIG. 14 is a diagrammatic view in section of a the nitrogen flow control unit of the fuel pressurization system;

FIG. 15 is a diagrammatic illustration showing the inter connected relationship entire hydraulic system;

FIG. 16 is a fragmentary elevation view showing the packaging arrangement of those hydraulic system components disposed between the wings of the missile;

FIG. 17 is a fragmentary elevation view showing the packaging arrangement of those hydraulic system components disposed between the wings of the missile and rotated about the missile axis 90.degree. from the position shown in FIG. 16, with the rotation being from bottom to top, it being noted that the turbine air intake, which appears midway between the wings is on the top centerline of the missile when the missile is in normal flight position;

FIg. 18 is a fragmentary elevation view showing the packaging arrangement of those hydraulic system components disposed between the wings rotated bottom to top about the missile axis 90.degree. from the position shown in FIG. 17;

FIG. 19 is a fragmentary elevation view showing the packaging arrangement of those hydraulic system components disposed between the wings and with the plane of view rotated bottom to top about the missile axis through an angle 90.degree. from the position shown in FIG. 18;

FIG. 20 is a fragmentary view in elevation with portions thereof broken away and shown in cross section as taken through the missile along line 20--20 of FIG. 18 and showing the arrangement of the hydraulic wing actuators;

FIG. 21 is a fragmentary view with portions thereof in section as taken on line 21--21 of FIg. 20 and showing the manner in which a hydraulic wing actuator is secured to the wing mounting bulkhead and one of the wing spars of a missile;

FIG. 22 is a fragmentary view with portions thereof broken away and in section illustrating, on a greatly enlarged scale, the main wing quick attach device;

FIG. 23 is a fragmentary view with portions thereof in section as taken along line 23--23 of FIG. 22;

FIG. 24 is a fragmentary view providing a showing similar to that of FIG. 23, with the exception that the latch levers and latch lever catches for wing attachment purposes are shown in a released position which would permit removal of the wing;

FIG. 25 is somewhat similar to FIg. 22 and shows the wing lock installation;

FIG. 26 is an axial section through the missile fuel tank and fuel pressurization compartment, many of the elements being shown in elevation;

FIG. 27 is a cross section taken on line 27--27 of FIg. 26 looking in the direction of the arrows and shows the packaging arrangement of the various fuel pressurization system components;

FIG. 28 is a fragmentary cross section taken on line 28--28 of FIG. 26 looking in the direction of the arrows, showing the bladder filling ports;

FIG. 29 is a fragmentary cross section taken on line 29--29 of FIG. 26 looking in the direction of the arrows, showing the fuel collection area;

FIG. 30 is a further section through the fuel collection area and is taken along line 30--30 of FIG. 29;

FIG. 31 is an axial section through the entire combustor section of the missile with many of the components being shown in elevation;

FIG. 32 is a perspective of the pilot fuel distributor, the secondary fuel distributor, and the shroud arranged in an assembled position, certain portions being shown in section or broken away;

FIG. 33 is a perspective of the combustor pilot assembly, certain portions being broken away;

FIG. 34 is a perspective of the combustor central can assembly with portions broken away;

FIG. 35 is an axial section through the secondary fuel distributor valve;

FIG. 36 is a perspective of the combustor outer shroud assembly with a portion broken away;

FIG. 37 is an elevation of the tailpipe assembly, shown as having been rotated 180.degree. from normal about the missile axis, with a portion being shown in section and further showing the arrangement of the coaxial cables and connections to the telemetering notch antennae;

FIG. 38 is a section through the missile tailpipe taken on the centerline of one of the fin supports, with the fin being shown in elevation;

FIG. 39 is a fragmentary cross section through the fin latching device taken on line 39--39 of FIG. 38, looking in the direction of the arrows;

FIG. 40 is a fragmentary cross section taken on line 40--40 of FIg. 38, showing in detail the connecting means between the coaxial cables of the missile body and a fin;

FIG. 41 is a detailed fragmentary elevation of the telemetering notch antenna on a greatly enlarged scale with the outer fin skin and antenna cover plate removed from the near side;

FIG. 42 is still a further enlarged cross section taken on line 42--42 of FIG. 41 through the telemetering notch antenna;

FIGS. 43A, 43B, 43C and 43D comprise a single erection flow diagram showing the manner in which all the various missile sub-assemblies are brought together to form the complete ramjet missile assembly which appears at the bottom of FIG. 43D, with all of the above assemblies being shown in perspective;

FIG. 44 is a side elevation of the missile and booster connected together, in an "on launcher" position;

FIG. 45 is a rear elevational view of FIG. 44;

FIG. 46 is a cross section on an enlarged scale, taken on line 46--46 of FIG. 44 looking in the direction of the arrows, through the missile aft clamping ring;

FIG. 47 is a longitudinal section taken on line 47--47 of FIG. 46 through the aft end of the missile and the forward end of the booster with certain elements being shown in elevation;

FIG. 48 is a fragmentary front elevational detail view, on an enlarged scale, of the missile-booster clamp ring release mechanism;

FIG. 49 is a longitudinal section taken on line 49--49 of FIG. 48 looking in the direction of the arrows;

FIG. 50 is a plan view of the mechanism shown in FIGS. 48 and 49;

FIG. 51 is a fragmentary longitudinal section, on an enlarged scale taken on line 51--51 of FIG. 46 looking in the direction of the arrows through the missile aft ring and booster forward head showing the means for indexing the booster to the missile;

FIG. 52 is an elevation of the detail shown in FIG. 51;

FIG. 53 is a detailed longitudinal section, on an enlarged scale taken on line 53--53 of FIG. 46, looking in the direction of the arrows through the aft and forward end of the missile and booster, respectively, illustrating the booster-missile separation signal arrangement;

FIG. 54 is a fragmentary plan of the arrangement shown in FIG. 53;

FIG. 55 is a fragmentary cross section taken on line 55--55 of FIG. 53;

FIG. 56 is a fragmentary longitudinal section, on an enlarged scale, through the aft and forward section of the missile and booster, respectively, illustrating the means employed to insure concentric spreading of the clamping ring upon missile-booster separation, and is taken on line 56--56 of FIG. 46 looking in the direction of the arrows;

FIG. 57 is a fragmentary cross section taken on line 57--57 of FIG. 56 looking in the direction of the arrows through the detail illustrated in FIG. 56;

FIG. 58 is a sectional view taken on line 58--58 of FIG. 56 looking in the direction of the arrows further illustrating the details shown in FIGS. 56 and 57;

FIG. 59 is an axial cross section taken on line 59--59 of FIG. 47 looking in the direction of the arrows through the forward end of booster;

FIG. 60 is a longitudinal section through the booster shroud, with the nozzle portions thereof being shown in elevation;

FIG. 61 is a side elevation, on an enlarged scale, of the launcher;

FIG. 62 is a detailed fragmentary elevation of the launching shoes and track;

FIG. 63 is a plan of the arrangement shown in FIG. 62;

FIG. 64 is an axial section taken on line 64--64 of FIG. 62 looking in the direction of the arrows;

FIG. 65 is a cross section, on an enlarged scale, taken on line 65--65 of FIG. 62 illustrating the details of the booster igniter electrical contact;

FIG. 66 is a cross section, partially in elevation, of the warhead;

FIG. 67 is a side view of the rod assembly in a partially extended condition;

FIG. 68 is a pictorial perspective illustrating the manner in which the rods of FIG. 67 expand in semi-circular ring sections upon detonation of the warhead;

FIG. 69 is a section through the nitrogen relief valve of the fuel pressurization system;

FIG. 70 is a section through one of the squib activated time delay units of the fuel system, certain elements being shown in elevation;

FIG. 71 is a section through the probe static relief valve of the fuel system;

FIG. 72 is a section through one of the flowmeters of the fuel system;

FIG. 73 is a section through the temperature pick-up of the main regulator of the fuel system;

FIG. 74 is a section through one of the pilot fuel injectors of the fuel system;

FIG. 75 is a section through one of the main injector nozzles of the fuel system;

FIG. 76 is an axial section through turbine and reduction gear assembly of the hydraulic system;

FIG. 77 is an axial section through the pump of the hydraulic system taken on line 77--77 of FIG. 78 looking in the direction of the arrows;

FIG. 78 is an end elevation of the hydraulic system pump illustrating the various pressure connections;

FIG. 79 is a detail elevation of the port plate in the hydraulic system pump illustrated in FIG. 77;

FIG. 80 is a section through the turbine control assembly of the hydraulic system;

FIG. 81 is a fragmentary section taken on line 81--81 of FIG. 80 looking in the direction of the arrows;

FIG. 82 is an axial section through the pump of the hydraulic system;

FIG. 83 is a semi-schematic axial section through one of the transfer valves of the hydraulic system;

FIG. 84 is a simplified schematic section through one of the torque motors of the hydraulic system;

FIG. 85 is a simplified axial section through one of the wing actuators of the hydraulic system;

FIG. 86 is an axial section through one of the wing locks of the hydraulic system;

FIG. 87 is a detail elevation of the piston and rod in the wing actuator illustrated in FIG. 86;

FIG. 88 is an axial section through one of the check valves of the hydraulic system;

FIG. 89 is an axial section through one of the filters of the hydraulic system, certain elements being shown in elevation;

FIG. 90 is an axial section through a second type of filter used in the hydraulic system;

FIG. 91 is an axial section through the solenoid valve of the hydraulic system;

FIG. 92 is a longitudinal section of a ramjet;

FIG. 93 is a plot of pressure of the air versus distance of travel along the ramjet illustrated in FIG. 92;

FIG. 94 is a plot of velocity of the air versus distance of travel along the ramjet illustrated in FIG. 92;

FIG. 95 is a plot of temperature of the air versus distance of travel along the ramjet illustrated in FIG. 92;

FIG. 96 is a longitudinal section through forward part of a ramjet illustrating the pressure forces in the diffuser;

FIG. 97 illustrates plots of various trajectories of the missile;

FIG. 98 illustrates plots of chamber pressure as a function of time for the booster propellant;

FIG. 99 illustrates plots of thrust developed as a function of time for the booster;

FIG. 100 is a schematic of the ignition circuit for the booster illustrated on sheet 39;

FIG. 101 is a schematic of the safety and arming system for the missile;

FIG. 102 is a block diagram of the shipboard weapons system for the missile;

FIG. 103 is a diagrammatic perspective of a shipboard control system for the missile;

FIG. 104 is a fragmentary side elevation of a shipboard launching system;

FIG. 105 is a fragmentary front elevation of the shipboard launching system illustrated in FIG. 104; and

FIG. 106 illustrates graphical representations of velocity-range characterizations for jet-propelled missiles.

GENERAL SUMMARY OF THE MISSILE

In accordance with the invention, there is provided a supersonic, surface-to-air, guided missile, launched by a solid-fuel booster, propelled by a ramjet engine, and deriving guidance intelligence from a ground-based radar system.

This missile is guided to its target in two phases, that is, a midcourse radar beam riding phase which brings the missile into approximate alignment with the target; and then, a homing phase for final guidance to the target.

The missile is provided with four wings 90.degree. apart and four fixed fins arranged in line with the four wings. The incidence angles of the wings are independently controllable, the wings being rotable about shafts disposed to protrude from the edge thereof which is normally in adjacency to the body of the missile and which shaft extends along an axis normal to the line of flight of the missile. Oppositely disposed wings moving together effect guidance control, and when moving differentially effect roll control. One pair of wings steers in the "A" channel (up-left, down-right), while the other pair of wings steers in the "B" channel (up-right, down-left).

For launching of the missile, a booster is suspended on a launcher rail by shoes near the fore and aft ends of the booster. The missile is suspended as a cantilever beam by means of a clamp ring that joins the aft end of the missile to the forward end of the booster. Upon missile launching, the clamp ring is opened, and the missile-booster combination is held together by acceleration forces until the booster propellant is expended.

Propulsion for the missile occurs in two stages. A solid-fuel booster launches the missile and accelerates it to the approximate operating speed of the ramjet engine. The ramjet engine then takes over and maintains active propulsion of the missile for the remainder of the flight of the missile to the aerial target.

The main propulsion of the missile is by its ramjet engine. Fuel is supplied to the pilot and main stages of the combustor of the ramjet engine through separate injection and regulator systems from a common supply system.

Referring now to FIGS. 1, 2, 3, 4, 5, 6, 43A, 43B, 43C, and 43D of the drawings, there is illustrated, in general, a missile 100 comprising the present invention. It is a ramjet, propelled, supersonic, surface-to-air guided missile for interception and destruction of aircraft carrying guided missiles or other types of destructive missiles.

The airframe assembly for the missile 100 is comprised of a number of major assemblies for ease of manufacturing and assembly, namely a forward body assembly 102 (including a nose section 104), a central body assembly 106, a combustor assembly 108, an aft body assembly 110, wing assemblies 112, 113, 114, and 115, and fin assemblies 116, 117, 118, and 119.

The forward body assembly 102 includes an inner body 120 including an accumulator 122, homing antennae 124 and waveguide 125, a static pressure probe 126, a ram pressure probe 128, body static pressure taps 130, a safety and arming device 132, a guidance system compartment 134, and a cover 135.

The center body assembly 106 comprises a center body 136 and a fuel tank 138, four wing locks 140 and brackets 142, an aft diffuser body 144, four wing actuators 1382, 1383, 1384 and 1385, four wing bearing sleeves 341, a center compartment cover 150, an aft compartment cover 152, and by-pass fairings 154.

The combustor assembly 108 is comprised of a forward shroud assembly 156, an outer shroud assembly 158, a pilot assembly 160, a central can assembly 162, and concentric fuel lines 164 and 166 to pilot injectors 330 and main fuel injectors 386, respectively, while the aft body assembly 110 includes a tailpipe 168, four fin supports 170, and a convergent-divergent exit nozzle 172.

Each of the four wing assemblies 112, 113, 114, and 115 are provided with a quick attachment device 174, while the fin assemblies 116, 117, 118, and 119 are provided with a fin attachment device 176 for each fin, and an antenna 178 for the beam riding system. Fins 116 and 118 are further provided with flare holders 180.

The body nose section 104 of the missile 100 includes compartments for a missile fuze system 182, which includes eight fuze antennae 184, collector rings 186, fuze electronics packages 188, and a fuze compartment cover 190; a safety and arming system 132, and the missile homing antennae system 124. The inner body 120 of the missile 100 is located in the body nose section 104, as will be pointed out more in detail subsequently.

The forward body assembly 102 of the missile 100 has a compartment 192 provided therein for housing a missile warhead system 194, or an exercise head assembly 196 including an exercise head structure 198, two telemeters 200 and 202, a distribution panel chassis 204, telemetering equipment 206, and two covers 208 and 210, therefor. Compartment 134 is provided for housing a steering intelligence system 212, a guidance control system 214, and an electrical power system 216.

In the center body assembly 106 of the missile 100, there is contained a compartment 218 for a hydraulic system 220; and compartments 222, 224, and 226 for a fuel system comprising the fuel tank 138, a fuel pressurization system 228, and a fuel control system 230, respectively. Hydraulic system compartment 218 and fuel control system compartment 226 are both located between the bulkheads 308 and 310 of center body assembly 106. In addition, the center body assembly 106 has mounted thereon the wings 112, 113, 114, and 115 for steering the missile 100 along its trajectory to the target.

The aft body assembly 110 of the missile 100, as previously indicated, includes the tailpipe 168, with the four fixed fins 116, 117, 118, and 119 mounted thereon, and the convergent-divergent exit nozzle 172. The combustor assembly 108 is located within the rear portion of the center body assembly 106 and the forward portion of the aft body assembly 110.

The missile airframe structure, FIGS. 1, 3, 44, 45, and 46, is comprised of the missile body 101, the four wings 112, 113, 114, and 115, the four fixed fins 116, 117, 118, and 119; and all of the structural provisions for installation of missile equipment, including all the structurally integral missile equipment. In conjunction with the missile booster 500, the airframe includes all fixed and removable missile-booster handling fittings; missile-booster structural and electrical connections and clamp ring 4036; forward and aft launching shoes 4030 and 4005, respectively, for the booster 500; four fixed fins 256, 257, 258, and 259; booster body 501; the nozzle 4000; and forward and aft handling shoes 4032 and 4007, respectively, for the booster 500.

The missile airframe structure also includes the inner duct 226, which has an inlet 268, a diffuser 270, a combustion chamber 272, and the convergent-divergent exit nozzle 172 located therein.

The inner duct 266 extends throughout the length of the missile 100 and is defined by the duct wall 274. Referring to FIGS. 1, 4, and 7, for purposes of illustration this duct 266 is divided into several main sections, namely, a cowl lip 276, a short conical section 278, a cylindrical intermediate section 280, and intermediate divergent or conical section 282, a second larger cylindrical section 284, and the convergent-divergent exit nozzle section 172, proceeding from the forward to the aft end of the missile 100. The second larger cylindrical section of the duct 266 includes the combustion chamber 272, and the tailpipe 168, while the other sections of the duct include the various stages of the diffuser 270, as will be described subsequently.

The surrounding skin 286 of the missile 100 defines the outer wall of the missile. At the nose section 104 of the missile 100, the outer wall 286 converges along a curve towards the inner wall. The forward ends of the outer skin 286 and inner duct 266, as best seen in FIG. 7, are attached to annular flanges 288 and 290 by welding or by means of screws. These flanges are integral with the cowl lip 276. The outer surface 292 of the outer skin 286 of the missile 100 is a continuation of the outer surface 294 of the cowl lip 276, while the inner surface 296 of the duct wall 274 is a continuation of the inner surface 298 of the cowl lip 276.

The cowl lip 276 defines the circular inlet 268 of the duct 266. At the rear of the missile 100, FIG. 37, the convergent-divergent nozzle 172 is secured by welding near the rear of the cylindrical section 284 of the duct 266, and defines a circular outlet 300 for the missile duct 266. This outlet 300 is used as an escape for the gases, which are liberated by the burning of the fuel.

In FIG. 4, transverse annular rings or partitions 302, 304, 306, 308, 310, and 312 are shown as mounted between the outer wall surface of the duct 266 and the inner wall surface of the outer skin 286. The partitions or transverse bulkheads 302, 304, 306, 308, 310, and 312 define the previously-mentioned compartments for receiving and housing the various missile component systems, as previously pointed out.

The inner edge of the annular rings or partitions are preferably welded to the missile duct 266, while the outer skin 286 of the missile 100 is made of semiannular covers of lengths corresponding to the lengths between the bulkhead rings or partitions. The covers comprising the outer skin 286 of the missile 100 are preferably attached to the bulkhead rings or partitions by means of screws or the like. This makes it possible to have ready access to any missile component system, by simply removing the particular cover that encloses the system.

As previously indicated, wings 112, 113, 114, and 115, FIGS. 1, 2, 43B, and 43C are considered as part of the airframe structure. With the exception of the mounting details, the structure of each wing is substantially according to conventional aircraft practices. All four wings are identical in construction.

Each wing has a high aspect ratio and tip rake, which tends to minimize hinge moment variations with Mach number under various combinations of wing deflection and missile angle of attack with the tapered planform of each wing tending to minimize the bending moments. The biconvex cross sectional shape of the wings minimizes drag for a given strength, simplicity of manufacture, resistance to chordwise bending, and elimination of corners which can affect aerodynamic properties.

As illustrated in FIGS. 1, 2, 43B, and 43C, the four steering wings 112, 113, 114, and 115 are oriented in a cruciform configuration and are located substantially near the center of gravity of the missile 100. These wings 112, 113, 114, and 115 are positioned in line with fins 116, 117, 118, and 119, respectively. The wings and fins are located substantially 45.degree. to the transverse axes of the missile 100.

Each wing 112, 113, 114, and 115 includes a main spar 301 with a plurality of ribs 303 spaced spanwise, as seen in FIG. 43B. Ribs 303 have fore and aft edges connected by leading and trailing edges 305 and 307, respectively. The entire wing structure is covered by skin 309, FIG. 43B.

In FIGS. 22, 23, 24, and 25, there is shown a typical wing 112 which has, at its root or inboard end, fore and aft rib members 311 and 313, respectively. Each root rib member 311 and 313 has an offset central portion 314 and an enlarged end portion 315. The enlarged end portion 315 of each rib member 311 and 313 is received between two brackets 316 and 317 which are formed integrally with the main spar 301. Bolts 318 and 319 are used to secure the enlarged portion 315 and the central portion 314 of each rib member 311 and 313, respectively, to brackets 316 and 317 on wing spar 301.

As illustrated in FIGS. 22 and 25, a wing mounting bulkhead 310 is mounted between the outer missile skin 286 and the diffuser 270. This bulkhead 310 is designed to receive a wing retaining sleeve 341. The root end 343 of the spar 301 is inserted in the sleeve 341. Roller bearings 321 and 325 of the needle type are located between the bulkhead 310 and sleeve 341 in order that the sleeve can be rotated in the bulkhead 310.

A rectangular shaped slot 327 is provided in the bulkhead 310 for receiving a wing actuating crank 329 which fits about the wing retaining sleeve 341 and which is locked thereto by means of threaded taper pin screws 331. The crank 329 has its driven end 333 slotted for receiving the piston rod 1386 from one of the wing actuators 1382 through 1385 as illustrated in FIGS. 20 and 21. A bolt 1388 is utilized to secure the end of the piston rod 1386 to the wing crank 329.

Wedged shaped tongues 335 are formed integral with the outer end of the wing retaining sleeve 341, and are spaced from the outer skin 286 of the missile 100. These tongues 335 are utilized to secure tapered end 337 of latch levers 339 in position so as to lock or secure wing 112 to the body of the missile 100. Each tongue 335 has a tapered aperture 345 for receiving the tapered end 347 of bolts 318, which, are used to secure the root rib members 311 and 313 to the wing spar 301.

As shown in FIGS. 22, 23, and 24, the latch lever 339 has a latch lever pivot boss 349 formed integral therewith at its end 337. This pivot boss 349 is comprised of an enlarged cylindrical portion 351, which is located in an aperture provided in bracket 317 adjacent to the end 337 of the latch lever 339, a conical-cylindrical central portion 353, a flange portion 355 which is received in a cylindrical recess 357 in the root rib members 311 and 313, and a reduced cylindrical end portion 359 which is inserted in an aperture 361 provided in the root rib members 311 and 313 and the bracket 316.

This pivot boss 349 is biased outwardly by means of a leaf spring stack assembly 363 as shown in FIGS. 22, 23, and 24. The spring stack assemblies 363 are secured to each root rib member 311 or 313 by bolts 365 and 367. End 369 of the longest leaf of spring assembly 363 has a U-shaped notch 371 provided therein. The notched end 369 of the spring assembly 363 is inserted between the flange 355 and conical-cylindrical portion 353 of the pivot boss 349. When the end 337 of the latch lever 339 is placed between the wedged portion 335 of the wing retaining sleeve 341 and the outer skin 286 of the missile 100, the leaf spring assembly 363 biases the flange portion 355 of the pivot boss 349 upwardly to jam the tapered end 337 of latch lever 339 under the tongue 335 of sleeve 341, thus pulling the wing 112 inboard and firmly seating the tapered end 347 of bolts 318 in the aperture 345 effecting thereby an extremely secure mount of the wing 112 to the missile 100.

The free end of the latch lever 339 is locked in its closed position by a latch lever catch 373 which is pivotally mounted on a screw 375. Screw 375, in turn, is secured to the root rib member 311 or 313. The latch lever catch 373 has a flange portion 377 formed integral therewith. One end of the flange portion 377 forms a lip 379, and is arranged to engage a notch 381 formed in the free end of latch lever 339, while the other end 382 is offset to underlie a portion of the wing skin 309 when the latch lever 339 is in a latched position.

A spring 383 is utilized to bias the latch lever catch 373 outwardly in order to secure the latch lever 339 in a latched position. One end of the spring is inserted in an aperture 491 provided in a tongue portion 385 of the latch lever catch 373, while the other end is secured to a pin 387. This pin 387, in turn, is secured to the root rib 311 or 313.

As previously mentioned, when the wing 112 is assembled to the missile 100, end 337 of each latch lever 339 is wedged beneath the tapered tongues 335 of the wing retaining sleeve 341. The other (notched) end of each latch lever 339 is secured against rotation by the lip 379 on flange 377 and a latch stop 389 formed integrally on the latch lever catch 373. When it is desired to remove the wing, or wings from the missile 100, the flange 377 of the latch lever catch 373 is pressed inwardly against the spring 383, FIG. 24, thereby rotating the lip 379 out of engagement with notch 381, and simultaneously bringing the tongue 385 into contact with the back side of latch lever 339 to kick the latch lever into the "unlatched" position.

As illustrated in FIGS. 15 and 25, each wing, such as wing 112, is provided with a wing lock 140 to prevent premature rotation of the wings. The winglock 140 is mounted on a winglock mounting beam or bracket 142. This beam is located between the bulkheads 308 and 310, and has its ends secured to clip members 391, which, in turn, are affixed to the bulkheads 308 and 310 by rivets 393.

Each winglock 140 comprises a housing having located therein a piston (not shown) which is actuated by hydraulic fluid applied through the tube 395. The piston drives a rod 397, one end 399 of which is received in a recess 401 provided in rib member 311 of the wing 112 to lock the wing against rotation. When it is desired to release the wing 112 hydraulic pressure is applied to the piston through tube 395 to retract the pin 397 from the recess 401 in member 311. The winglock 140 can be manually operated by inserting the end of a suitable tool through openings 403 and 405 provided in the missile outer skin 286 and beam 142, respectively, to engage a notch 407 provided in a release latch 409. The release latch 409 is provided with pawl 411 which engages a notch 413 in rod 397. The release latch 409 is pivotally mounted, by pin 415, to two brackets 416 which, in turn, are integrally secured to the winglock 140. A spring 417, having one end thereof connected to pin 415 and its other end resting against a pin 418 mounted on the release catch 409, is utilized to bias the release latch 409. By applying a force, in a clockwise direction, to the release latch 409, it is pivoted in a clockwise direction about the pivot pin 415 to withdraw the rod 397 from the recess 401 in rib 311 to release the wing 112.

All four tail fins 116, 117, 118 and 119 are alike with the exception that fins 116 and 118 carry the telemetering notch antennae 178, and fins 117 and 119 are provided with tracking flare holders 180.

As shown in FIG. 38, each tail fin 116, 117, 118, and 119 is constructed according to conventional practices and is comprised of leading and trailing edge members 419 and 421, respectively, two spars 423 and 425, a plurality of rib members 427 spaced spanwise of the fin, doublers 429, and fore and aft root end ribs 431 and 433, respectively. Each tail fin has an identical biconvex or double circular arc cross section for the same reasons as the wings 112, 113, 114, and 115. The entire tail fin structure is covered by skin 435. The mounting structures for the four fins are identical in construction.

As illustrated in FIGS. 37, 38, and 39, four mounting rings 437 are provided between the wall 284 of the tailpipe 168 and a fairing 439 for supporting the four fin supports 170. These fin supports 170 are utilized for attaching the tail fins to the missile 100. Each fin support 170 has four arc shaped members 441, FIG. 39 formed integral therewith, two at each end thereof for securing the fin support 170 to the fin mounting rings, by suitable means, such as rivets 443 and bolts 445.

Each fin support 170, is comprised of a body which is rectangular in cross section and which is arranged to be inserted into a rectangular shaped recess 447 provided in the root end of each fin. Holes 451 and 453 are provided in the fin support body for receiving pins 455 and 457, respectively, which are formed integrally with the fin attachment device 176. The fin attachment device 176 is sandwiched between the doublers 429 and the skin 435 of the fin. The fin attachment device is arranged in the fin so as to allow the pins 455 and 457 to protrude from the base of the rectangular shaped recess 447 in the fin.

The body of each fin support 170 has an aperture 459 provided therein for receiving a latching mechanism 461 which is utilized to secure the fin 116 to the fin support 170 on the missile. The aperture 459 is counterbored at 463 to receive a coil spring 465.

A latching piston 467 is located in the aperture 459. This piston 467 has a core 469 with a lip 471 formed therein. The coil spring 465, in counterbore 463, is used to bias the piston 467 into a locked position.

A plate 473 is secured to one end of the piston 467 by a screw 475. This plate 473 maintains the piston 467 in the aperture 459 when the fin 116 is not secured to the fin support 170. This plate 473 is received in a small recess 477 in the fin support 170.

The fin attachment device or fitting 176 carries between its two pins 455 and 457 a headed lug 479 which cooperates with the lip 471 in core 469 of the piston 467 to securely lock the fin 116 in place after installation on the missile 100. The head 481 of the lug 479 is tapered to facilitate its insertion into the core 469 of piston 467. The piston 467 is held against rotation in the aperture 459 by means of a pin 483 press fitted into a hole in the fin support 170. This pin 483 slidably engages a slot 485 milled in the surface of the piston 467.

In operation, when the fin is assembled to the fin support 170, the pins 455 and 457 are inserted in their holes 451 and 453, respectively, and the lug 479 is inserted into the fin support 170 through an aperture 487 provided in the fin support. The tapered head 481 of the lug 479 rides against the sloping face 489 of the lip 471, forcing the piston 467 to the left, until the head 481 of lug 479 enters the core 469 allowing the spring 465 to pull the lip 471 into locking engagement under the head 481 of lug 479.

The fin can be removed from the fin support 170 by pushing the piston 467 to the left to withdraw the lip 471 from beneath the head 481 of the lug 479, and then withdrawing the pins 455 and 457, and lug 479 from their cooperating apertures in the body of the fin support 170.

The overall propulsion system for the missile 100 includes the booster 500 for projecting the missile to the desired separation velocity for operation of the ramjet engine, and the ramjet engine for accelerating the missile 100 from the booster-missile separation velocity to its design velocity where the velocity of the missile 100 is then stabilized by the missile fuel control system 230. The propulsion system includes the aforementioned missile fuel system including the fuel tank 138, the fuel pressurization system 228, and the fuel control system 230.

The ramjet engine for the missile 100 is designed to accelerate the missile 100 from the velocity that the missile 100 has at the time of its separation from the booster rocket 500 to its design velocity. The velocity of the missile 100 is then controlled by the fuel control system 230, as previously pointed out. In general, the ramjet engine includes the single oblique shock, spike diffuser 270, the piloted can combustor 108 and the convergent-divergent exit nozzle 172.

In this missile 100, at the design speed, the supersonic air stream passes through the oblique shock wave which is centered at the nose of the inner body 120 of the missile 100. In passing through this shock wave, the velocity of the air stream is decreased to a lower supersonic velocity. At the entrance 268 of the diffuser 270, the air stream passes through a normal shock wave and its velocity is then reduced in the first stage (short conical section) 278 of the subsonic diffuser to a high subsonic velocity. The air is then expanded in the second stage (intermediate cylindrical section) 280 of the subsonic diffuser 270 to a lower subsonic velocity. Since an increase in pressure in the diffuser 270 is associated with a decrease in velocity therein, the pressure in the diffuser 270 is increased. The air is then mixed with the fuel and is burned continuously in the combustor 108. This maintains the high pressure in the diffuser 270 and also increases the volume of the air by increasing its temperature in the combustor 108.

The thrust developed by the missile ram engine is, therefore, equal in magnitude to the difference of momentum between the incoming air and the outgoing products of combustion as best seen from Equation (2) below:

T = [(M + m)v.sub.2 - Mv.sub.1 ] Eq. (2).

where

T = thrust force in pounds;

M = mass flow of air in slugs per second;

m = mass flow of fuel in slugs per second;

v.sub.1 = incoming velocity of air in feet per second; and

v.sub.2 = outgoing velocity of products of combustion in feet per second.

FIG. 92 illustrates diagrammatically the ramjet engine for the missile 100. It has a multiple shock diffuser 270 to slow the incoming air and convert the velocity energy to pressure energy. It is composed of the inner body 120 and the aerodynamic (inner) duct 266. The multiple shock diffuser 270, with the projecting inner body 120 forming shock waves at the diffuser entrance, reduces the velocity in easy stages through one or more angle shock steps before it reaches the final weak normal shock that drops the velocity to a subsonic value. The normal shock, in this case is much less severe than one across an open duct. Using this method, therefore, more ram pressure can be recovered and put to use.

Assuming that the ramjet engine has been boosted to a Mach number M = 1.7, a cycle through the engine will be followed by reference to FIGS. 92, 93, 94 and 95. At point 1, ambient conditions exist. At point 2, the first angle shock is formed by the projecting inner body 120. It has been found that as air flows through a shock wave, the pressure increases abruptly, the velocity decreases, and the temperature increases. Point 2 on the three curves in FIGS. 93, 94 and 95 shows these changes. Between points 2 and 3, the air has been slowed to a velocity somewhat less than M = 1.7. At point 3, a second, weaker angle shock wave is formed. As the air passes through the shock wave, the same type of changes take place (see point 3 on the curves). Between points 3 and 4, the air has a velocity less than between points 2 and 3, but still greater than M = 1.0. At point 4, the weak normal shock wave is formed. This also causes the same type of changes and finally drops the airflow to subsonic conditions. From this point on, the airflow greatly benefits from the law of continuity (the mass flow (.rho. AV) = a constant) and Bernoulli's theorem (as the velocity decreases, the pressure increases). By the law of continuity, as the area of the diffuser 270 increases, the velocity decreases to keep the mass flow constant. Since the velocity constantly decreases, the pressure increases.

Leaving the diffuser 270, the air flows into the combustor chamber 272. The correct balance between the combustion chamber pressure and the diffuser exit pressure is quite critical for the proper operation of the ramjet engine. If the combustion chamber pressure is too high, the diffuser pressure will be too high and will tend to push all the shock waves upstream slightly, detaching the first angle shock wave. This will cause "spillover" which effectively diverts some of the incoming air that is needed for the proper operation of the engine. If the combustion chamber pressure is too low, the diffuser pressure will also be too low and the shock waves will all be moved downstream slightly, the first angle shock wave being swallowed. With this condition existing, large pressure losses, turbulence, and heat generation occur, thereby lowering the efficiency of the engine.

Assuming now that the correct combustion chamber pressure exists, the following action takes place. The air from the diffuser 270 is mixed with fuel, the mixture is ignited and burned. The increases in temperature increases the volume of the air mass. The hot expanding gases try to find an exit, but are blocked radially by the walls and upstream by the diffuser pressure. Because of this flow restriction, the diffuser pressure increases and forces the combustion gases out the rear exit at a very high velocity. The increased diffuser pressure is transmitted upstream to the incoming air. This action allows the incoming gases to remain compressed, for without this back pressure the pressure in the diffuser would not be maintained.

The ramjet missile 100 is pushed along in aerial flight by two natural occurrences, namely the high diffuser pressure created by the forward motion and that of the hot gases formed by combustion; and the fact that pressure is exerted equally in all directions and acts perpendicular to a surface.

As shown in FIG. 96, pressure acts normal to the diffuser walls and the inner body 120. However, each pressure force can be broken into two force components, one acting laterally (perpendicular to the engine axis) and one acting longitudinally (parallel to the engine axis). All the lateral force components cancel out, eliminating any unbalance in the lateral direction. All the longitudinal force components act in the same direction, and when summed up, equal the output thrust from the ramjet engine, a force which is large enough to stabilize or accelerate the engine at the designed supersonic speeds.

The chosen trajectory for the missile 100 is determined by the altitude and range of the target with the consumption of fuel being the criteria for the chosen path of travel of the missile 100. At high altitudes, the drag and air mass intake to the ramjet engine are low. Since the thrust developed by the ramjet engine is dependent upon the air-fuel ratio in the combustor 108, low fuel flow rates are needed for aerial flight of the missile 100 at high altitudes. For long range flights, an up-and-over trajectory will be used for the missile 100. For shorter ranges, where fuel consumption is not critical and tie-up time of the radars is more important, a more direct path of flight of the missile 100 will be used. This is illustrated in FIG. 97.

The multi-shock diffuser 270 is utilized to slow the incoming air and to convert its velocity energy to pressure energy. Fuel is then injected into the airstream passing through the missile 100, and the air-fuel mixture is then introduced into the combustor 108.

This combustor 108 is composed of two stages; namely, the pilot stage and the main stage. Each stage has its own fuel distribution system which supplies it with the proper air-fuel mixture to maintain combustion in the combustor 108. The gases liberated due to the burning of the fuel escape through the exit nozzle 172.

The diffuser assembly is comprised of a Ferri type inlet designed for high pressure recovery with low additive drag over the design Mach number range, a cylindrical section, and a conical-cylindrical section which encloses the combustor pilot 320, FIGS. 1 and 7.

The diffuser assembly, in particular, comprises a single cone, Ferri type supersonic intake, which is followed by a two stage subsonic diffuser 270. As indicated in FIG. 7, this intake is comprised of the circular inlet 268 in the center of which is placed the streamlined inner body 120.

This inner body 120 is supported by three strut members 322, 323, and 324, which are attached at one end to the inner body 120, with the other ends thereof being attached in the inner wall of the conical section 278 of duct 266, FIGS. 2 and 7.

The distinctive feature of the inner body 120 is the conical nose 326, which has an angle of approximately 22.75.degree., causing a single conical oblique shock wave which accounts for the external supersonic compression of the air before it enters the inlet 268 of ramjet missile 100. In the vicinity of the cowl lip 276, depending on the particular configuration, a normal shock wave occurs. These two shock waves complete the supersonic compression of the air.

In addition to supersonic compression of the air, the air undergoes two stages of subsonic compression in the subsonic diffuser 270 which includes the first cylindrical section 280, the divergent or conical section 282, and the second cylindrical section 284, of the missile duct 266. The first stage of subsonic compression of the air occurs in the duct between the cowl lip 276 and the exit 328 of the first cylindrical section 280 of the duct 266, while the second stage of subsonic compression of the air occurs in the divergent section 282 and the second cylindrical section 284 of the duct 266.

The combustor 108 of the ramjet engine is that part of the ramjet engine in which the fuel is injected and burned. One of the characteristics of the combustor 108 is that it delivers the heat value contained in the injected fuel to the air flowing through the combustor 108 with a high degree of efficiency and low aerodynamic drag. The combustor 108 is also capable of sustaining burning over a wide range of air flows. The combustor 108 was designed to occupy a minimum of space and a small fraction of the total weight of the missile 100.

The speed and altitude requirements of the missile 100 determines that the combustor 108 be required to operate at pressures as low as 5 psia pressure (high altitude requirement), and that a minimum cruise combustion efficiency of 80 percent at high altitudes and 90 percent at low altitudes be obtained. Cruise thrust requirements for the missile 100 demanded fuel-to-air equivalence ratios ranging from 0.2 to 0.6 be used.

Referring now to FIGS. 31 through 36, the pilot section 320 of the combustor 108 contains pilot fuel injectors 330, pilot shrouds 332 and 334, and an annular pilot combustor can or flameholder 336. It is the function of this pilot section 320 to ignite and stabilize the burning of the main fuel and air mixture in the combustor 108. Fuel for the pilot injectors 330 travels down a centrally located tube 164 from a pilot regulator 340 to four radially mounted pipes 342 which carry the fuel to the pilot fuel manifold 344 carried between the shrouds 332 and 334 forward of the pilot flameholder 336. Air for the pilot section 320, approximately 20 percent of air flow through the ramjet engine, is captured by an annulus 346 formed by the inner and outer pilot shrouds 332 and 334, respectively, and located in the divergent section 282 of the diffuser duct 266. The fuel is injected into the air stream from the manifold 344 by sixteen spring loaded pintle type injectors 330 which will be described later. Fuel is injected forward into the air stream (that is, in a direction contra to the direction of movement of the air stream) in a fine spray to insure thorough mixing with the air and rapid vaporization thereof.

As illustrated in FIGS. 31 and 33, the air-fuel mixture then passes through apertures 348 in the walls 350 and 352 of the stainless steel pilot flameholder 336 to the inside of the pilot flameholder 336. Once inside the pilot flameholder 336, the air-fuel mixture burns by the flame holding on the incoming air jets. The velocity of the air-fuel mixture entering the apertures 348 is such that flashback and burning of the pilot walls 350 and 352 is prevented. Cooling louvers 354 are provided between the apertures 348 to prevent these areas from becoming overheated. The flame from the pilot flameholder 336 exhausts to the main burning area without restriction. At its forward end, the pilot flameholder 336 is supported in the combustor 108 by four slotted blocks 356 which cooperate with lugs 358 which are welded to struts 360. Struts 360 are, in turn, bolted to four combustor supporting stringers 362 which are, in turn, riveted to the intermediate conical section 282 of the duct 266. In order to provide for expansion and contraction of the parts due to heating and cooling, the after end of the pilot flameholder 336 is loosely supported by means of flanged rings 364 and 366. Flanged ring 364 engages a channeled ring 368 on the combustor central can assembly 370 and ring 366 engages a cylindrical tongue 372 on the combustor cooling shroud assembly 374.

The main combustor 376 for the missile 100 is comprised of a main fuel injector 378, the combustor central can assembly 370, the combustor cooling shroud assembly 374, and the convergent-divergent exit nozzle 172. The fuel for the main combustor 376 is transferred from a main fuel regulator 380 by the main fuel delivery tube 166, which is supported concentrically within the pilot fuel delivery tube 164, to a main fuel distribution valve 384 located at the forward end of the combustor central can assembly 370. This valve 384 is a spring loaded sliding piston valve that has a maximum pressure drop of 30 psi. This fuel distribution valve 384 divides the fuel flow into sixteen equal parts and ditributes this fuel to each of sixteen main fuel injectors 386, which will be described later.

These injectors 386 are equally spaced around the inside of the combustor central can assembly 370. The forward end 388 of the central can assembly 370 is open to free flow of air from the diffuser 270, and is of such size as to capture 60 percent of the air entering the combustor 108.

The air, after mixing with the fuel, flows radially outward through the apertures 390 in the combustor central can assembly 370 into the main burning area, which is merely as extension of the exit of pilot flameholder 336. All burning of the main air-fuel mixture takes place after passing through the apertures 390 in the combustor central can assembly 370 with the flame holding on the air jets entering the burning zone. The apertures 390 in the combustor central can assembly 370 provide sufficient cooling with the rsulst that louvers are unnecessary. The combustor central can assembly 370 is supported by four stringer-type hangers 392. Hangers 392 are provided at their forward end with studs 394 which fasten to the four main support struts 360 of the combustor 108.

As shown in FIGS. 31 and 36, the combustor cooling shroud assembly 374 is located externally with respect to the burning zone in order to keep the tailpipe 168 cool by channeling air from the diffuser 270 between the tailpipe 168 and the combustor shroud assembly 374. Approximately 20 percent of the air flow through the ramjet engine is used for cooling. The shroud assembly 374 itself is cooled by two rows of small apertures 396 and 398 at its forward end 400 and four rows of louvers 402 which direct relatively cool air over its inner surface 404. This air does not have any fuel injected into it. The air, after passing through the cooling shroud assembly 374 through apertures 396 and 398, and louvers 402, mixes with the main air stream and aids in burning of the air-fuel mixture. A gap 406, provided between the shroud 374 and the tailpipe 168, permits a flow of cool air to pass over the exit surface of the nozzle 172, thus cooling it to prevent burn out thereof.

In FIGS. 31 and 36, the forward section 400 of the shroud 374 is supported in the combustor 108 by four pairs of lugs 408 which bolt to struts 360, which are, in turn, bolted to stringers 362 in the combustor 108. The aft section of the shroud assembly 374 is supported in the combustor 108 by a series of thin flat strutlike members 410 which are formed from continuous strips and welded to the outer surface of the shroud assembly 374. The outermost webs 412 of struts 410 are capped by a strap 414 encircling the entire shroud assembly 374.

Due to the limited space available for the combustor 108 in the missile 100, it is necessary to insert the combustor pilot section 320 as far as possible into the exit of the diffuser 270. Inserting the pilot section 320 into the diffuser exit aids in obtaining high combustion efficiencies.

A high degree of burning stability with high combustion efficiency has been achieved in the missile 100 by the use of a pilot section 320 with a large heat release (approximately 20 percent of the total air flow through the missile 100) and the use of a nearly constant air-fuel ratio in the pilot stage of the combustor 108. The durability of the combustor 108 is insured by the use of the cooling louvers 354 and 402, and the shroud assemblies 332, 334, 336, 370, and 374.

At separation of the missile 100 and the booster rocket 500, where the air loads on the combustor 108 are high, supersonic air flows (hence low pressures) may be present in parts of the combustor 108, with high differential pressures across components. Since these parts have not been heated by burning, their strength will be a maximum and no damage will result.

The initial thrust for the missile 100 is supplied by the solid propellant booster rocket 500 which separates automatically at the end of the boost phase. At separation of the booster rocket 500 and missile 100, the missile will be travelling at a Mach number where, after rapid ignition of the fuel takes place, sufficient thrust for acceleration will be produced by the ramjet engine. The missile 100 will accelerate to design velocity and then maintain this velocity through the remainder of the flight to an aerial target. To accomplish this, the missile 100 carries a fuel supply 138, a fuel pressurization system 228, fuel injectors 320 and 378, a fuel control system 230, and a combustor 108 necessary for operation as a ramjet engine.

In FIGS. 8, 9A, 9B, 26, 27, 28, and 29, there are shown the working components of the complete fuel system. The fuel control system 230 begins with the supply of the fuel from the fuel tank 138 and ends with the delivery of the correct amount of fuel to the pilot and main fuel ducts 420 and 422, in an air intake strut 424, which carry the fuel back to the pilot and main injectors 320 and 378, respectively, of the combustor 108.

The ramjet engine fuel is contained in an annular tank 138 which is provided with a quickdisconnect fuel filler valve 426. An outlet line runs from this tank 138 to a "G" operated fuel shutoff valve 428 which holds the fuel in the tank 138 until the booster rocket 500 operation provides the necessary inertia force to trip a valve opening mechanism 430. Provision is made for expansion of the fuel during storage in the fuel tank 138.

Fuel is forced from the tank 138 and kept under pressure by means of two pressurized flexible bladders 438, located in the tank 138. This insures delivery of fuel to a fuel pump 440 at pressures sufficient to prevent cavitation of the pump under all flight conditions. The pressurization of the bladders 438 is obtained from a tank 442 of high pressure nitrogen and is controlled by a two stage nitrogen flow control or regulator unit 444. This unit 444 also contains a boost inertia force tripping G valve 446 and a surge-preventative piston valve 448.

In low altitude flight, the nitrogen supply is rapidly used up and an alternate pressurizing source incorporating pressurized air from the diffuser 270 is utilized to pressurize the fuel in the tank 138. The alternate pressure source is utilized when the regulated nitrogen pressure drops below the inlet air pressure of the turbine pump 440 (diffuser pressure minus scoop losses), whereupon a check valve 450 opens and the diffuser air pressure then supplies the pressurization for the bladders 438 in the fuel tank 138.

A safety relief valve 452 is incorporated in the nitrogen pressurizing system and referenced to the pressure in the diffuser 270 to limit the pressure across the inner wall 454 of the fuel tank 138, which is the limiting portion of the structure of the fuel tank 138.

The turbine driven fuel pump 440 is utilized for pressurization of the fuel. This turbine is a single stage impulse type with supersonic nozzles 456. The turbine fuel pump 440 is of the single stage centrifugal type. The turbine fuel pump 440 maintains an outlet pressure above the minimum required for fuel flow control. This pressure is dissipated in line losses, strainer loss, regulator drop and injector drop, plus diffuser pressure. The pilot and main fuel regulators 340 and 380, respectively, require a certain throttling pressure drop to operate and are capable of throttling considerably more than this. The excess pump outlet pressure is then throttled by these regulators 340 and 380, with the result that a reserve of pressurized fuel is thus provided to take care of varying demands and also to provide a margin of safety.

The sliding parts of the fuel regulators 340 and 380, and injectors 330 and 386 are of precision construction with clearances requiring dirt free operation. A fuel strainer 458 removes all contaminating materials of sufficient size to cause binding of the sliding parts. As the fuel leaves the strainer 458, it is divided into two streams to supply clean fuel to both the pilot and main regulators 340 and 380.

The pilot regulator 340 controls the fuel flow rate to the pilot section 320 of the combustor 108, maintaining a nearly constant air-fuel ratio which insures stable combustion at all times. The main regulator 380, on the other hand, controls the fuel flow rate to the main section 376 of the combustor 108. The main regulator 380 incorporates a velocity control mechanism which causes a shift in fuel flow from a rich limit to a lean limit in the design velocity range. The rich limit operation will produce sufficient thrust to give acceleration of the missile 100, while lean limit operation, which will be nearly zero flow through the main regulator 380, will result in deceleration of the missile 100. This will give a constant velocity of the missile 100 at some flow between the rich and the lean limits.

Fuel flowmeters 460 and 462 are provided at the outlets of both regulators 340 and 380, respectively, for measurement of fuel flow through the pilot and main systems.

The fuel storage tank 128, is annular in shape and contains the supply of fuel, preferably kerosene, necessary for operation of the ramjet engine of the missile 100.

The fuel storage tank 138 is located in the fuel tank compartment 222 which occupies a section of the missile 100 beginning at approximately mid-chord of wings 112, 113, 114, and 115. It is an integral part of the missile airframe, and is formed by the outer surface 464 of the diffuser of duct wall 274 reinforced by five stiffening rings 466 and the inner surface of the missile skin 286. Both the missile skin 286 and the diffuser wall 274 are made of stainless steel in this region. The ends of the tank 138 are formed by annular stainless steel rings or bulkheads 468 and 470, hemispherical in cross section, welded to the missile outer skin 286 and the diffuser wall 274. The tank is fitted with a fuel drain plug 472 on the bottom and has an air vet plug 474 provided in the top thereof.

As shown in FIGS. 26, 28, 29, and 30, the tank 138 contains two longitudinal partitions 476 and 478 to separate the two bladders 438. The partitions 476 and 478 are constructed so as to form perforated triangular longitudinal fuel feeder channels 480 and 482. Welded to the end walls 468 and 470 are perforated conical rings 484 and 486, respectively, which cooperate with said end walls to form annular fuel feeder channels 488 and 490. Two perforated gusset plates 492 are provided at the intersection of longitudinal feeder channel 482 and annular feeder channel 488 so as to form a fuel collection area 494. Thus, all fuel feeder channels are in communication with the fuel collection area 494 permitting fuel in any part of the tank 138 to reach the fuel outlet 436.

The tank 138 contains two access doors 496 for installation of the bladders 438. These doors 496 are secured in place by machine screws and sealed against leakage by an "O" ring around their inner surfaces. The forward bulkhead 468 of the tank 138 is fitted with two bosses 498 through which the fittings for the tank bladders 438 pass and connect with nitrogen pressurization lines 502.

Bladder fittings 504 are held in the tank boss 498 by lock nuts 506. An effective nitrogen seal at the bladder fittings 504 is achieved by clamping the bladder 438 between the head 508 of the fitting 504 and a washer 510. A conventional gasket 512 between the tank boss 498 and the washer 510 provides a simple fuel seal. Other bosses 514 and 516, as shown in FIG. 9B, in the forward bulkhead 468 are provided for connecting the fuel line 436 to the fuel shut-off valve 428 and to vent the fuel line 436 to the inlet side of the fuel shutoff valve 428 back to the tank 138.

The 2 1/4 inch aluminum fuel line 436 is connected to the tank boss 514 by a short flexible hose 518 which is secured by clamps 520. A boss 521 located in the aft tank bulkhead 470 connects the tank 138 to the fuel fill port valve 426 which is utilized to fill the tank 138 with fuel through an internal tube fitting (not shown). This port valve 426 is a quick disconnect fitting that is self-closing when the external fuel fill line (not shown) is disconnected. The tank 138 is an integral part of the missile airframe, and is a load bearing part of the missile 100 in aerial flight.

The missile fuel tank 138 contains two, two-ply nylon fabric, neoprene-impregnated bladders 438. Each bladder 438, when fully inflated, occupies approximately one-half of the missile tank volume. The bladders 438 are installed in the fuel tank 138 through the access doors 496 and are held securely in place by the bladder fittings 504 and by a plurality of snap fasteners 522 which lock in holes 524 that are provided in the longitudinal feeder channels 480 and 482. To facilitate easy installation and to minimize the number of openings on the bladders, each bladder 438 has bonded to it two longitudinal tongues 526, the free ends of which are provided with a series of snap fastener holes corresponding to the number of snap fastener holes 524 in the longitudinal feeder channels 480 and 482. The bladders 438 conform to the inside shape of the tank 138 when fully inflated, and are completely supported by the tank walls, ends, and partitions except for the spaces where the holes are drilled in the partitions and conical plates for fuel passage.

When the missile 100 is in fueling position, a vent line 432 is provided from the top of the fuel shut-off valve inlet 434 to the top section of the fuel tank 138 to eliminate any trapped air from a high spot which exists at the inlet of the fuel shut-off valve 428 and to insure complete fuel charging of the tank 138.

In FIGS. 9A and 9B, the fuel in the tank 138 is pressurized, from two sources, namely the inert gas tank 442 or, scoop 772 which extracts diffuser pressure from the missile duct 266 when the nitrogen pressure drops below the pressure in the diffuser 270. The diffuser pressure is then supplied to the bladder 438 through a line 528, check valve 450, a line 530, and line 502. Fuel pressurization is necessary in order to provide fuel at the inlet 532 of the fuel pump 440 at sufficient pressure to prevent cavitation of the pump, to provide sufficient fuel pressure at high altitudes, and to eliminate boiling of the fuel in the fuel tank 138. The minimum pump inlet pressure required to prevent cavitation of the pump 440 depends upon the volume of fuel flow, fuel vapor pressure (fuel temperature), and, to a small extent, missile accelerations.

When the fuel is pressurized by the inert gas pressurization system, which utilizes the gas, such as nitrogen, the gas is supplied from the nitrogen tank 442 which is provided with two hemispherical ends 534 and 536 and the body 538. End 534 of the tank 442 is fitted with a boss 540 for a clean-out plug which is sealed by an O ring. Near the opposite end 536 of the tank 442, a fitting 542 is affixed to the tank body 538. A tube 544 is connected to the fitting 542, and to the nitrogen flow unit 444.

The nitrogen tank 442 is filled by a tire type fill valve 546, as seen in FIGS. 9B and 27, which is attached to a fill and gauge manifold 548. This valve 546 incorporates a secondary metal-to-metal sealing seat to reduce leakage of the nitrogen. A pressure gauge 550 is mounted in the filler and gauge manifold 548 to give a visible indication of the charging pressure in the nitrogen tank 442.

The fuel control system, as indicated above, incorporates the fuel pressurization bladders 438 in the fuel tank 138. Each bladder 438, external of the tank 138, is, in turn, connected to the line 544 supplying the inflation nitrogen pressure from the nitrogen tank 442.

When fully extended, the bladders 438 are supported by the walls of the fuel tank 138, except for a considerable number of holes in the various members 476, 478, 484, and 486 in the fuel tank structure. The bladders 438 conform to the hat shape stiffener rings 466 in the fuel tank interior and are held in place longitudinally, by the hinge type flaps or tongues 526 which are fastened to the fuel feeder channels 480 and 482 with snap fasteners 522.

A quick disconnect bleeder valve 552 is provided in the pressurizing line 502 to vent the bladders 438 and permit them to collapse during the fueling process.

The nitrogen flow control unit 444, as best seen in detail in FIG. 14, is provided to supply the nitrogen at a pressure equal to ram pressure plus 28 psi to the fuel pressurization bladders 438 and to prevent leakage of high pressure nitrogen during storage thereof. When the pressure of the nitrogen dorps below that of the pressure of the diffuser 270, the check valve 450 opens and the fuel tank bladders 438 are then pressurized by the larger pressure of the diffuser 270.

It has been determined that the nitrogen must be supplied at the pressure indicated based on the fuel flows required in the missile 100 and the fuel tank size, minimum fuel pump inlet pressures to prevent cavitation, acceleration effects of the missile 100 on pump inlet pressures, and fuel vapor pressure variation with fuel temperature. The flow control unit 444 is sealed against nitrogen leakage by the use of O rings and leather gaskets. The pressure from the nitrogen tank is applied through tubing 544 to the inlet 562 of the surge preventing valve assembly 448.

The surge preventing valve assembly 448 is comprised of a sleeve 564 secured in the body casting 560 by a steel snap ring 566. Four ports 568 permit the nitrogen to flow from the center of the sleeve to an annulus 570 around the sleeve 564. This annulus 570 leads to the inlet 572 of the shut-off valve 446. Piston 449, which prevents surge of the nitrogen, fits inside of the sleeve 564. The inlet end 574 of the piston 449 has four V-shaped slots 576 cut into it. The opposite end 578 of the piston 449 is closed, and is fitted with a flange 580 to position it in the closed position.

A spring 582, compressed between a piston stop 584 and the top 578 of the piston 449, holds the piston 449 in the closed position when no nitrogen is flowing. The piston stop 584 is threaded for purposes of adjustment and is locked in place by screw 586. a boss 588 on the stop 584 limits the travel of the piston 449 and also maintains the spring 582 in a central position.

The cover 590 for the surge preventing valve assembly 448 is screwed into the main body casting 560 and is sealed by an O ring 592. The space between the top of the piston 578 and the piston stop 584 is connected to the shut-off valve inlet passage 572 by a drilled passage 594 fitted with a No. 80 bleed 596.

When the nitrogen tank 442 is charged, with the nitrogen shut-off valve 446 closed, nitrogen tank pressure is applied to both sides of the surge preventing piston 449 by leakage past the piston 449. When the nitrogen shut-off valve 446 is opened, a sudden decrease in pressure on the surge preventing valve outlet 598 causes the valve 448 to open.

However, the high pressure nitrogen on the back side of the piston 449, bleeding off through the No. 80 bleed 596, causes the valve 448 to open slowly, and prevents a surge of high pressure nitrogen from passing through the flow control unit 444. The spring 582 tending to close the piston 449 causes a small pressure drop to be generated across the surge preventing valve 448.

The outlet port 598 of the surge preventing valve 448 is directly connected to the inlet 572 of the nitrogen shut-off valve 446. This valve 446 is of the inertia activated type. It opens during launching of the missile 100. The shut-off valve 446 is locked in the closed position by six stainless steel balls 600 which ride in a groove 602 on the valve stem sleeve 604. When the valve is open, it is locked in the open position by the same steel balls 600.

The nitrogen shut-off valve 446 consists of a valve disc 606, a valve seat 608, a valve stem 610, the previously mentioned valve stem sleeve 604, a ball retaining sleeve 612, a G weight 614, a shut-off valve cover 616, two springs 618 and 620, the six steel balls, 600, and various O rings and gaskets. The six steel balls 600 fit into holes 622 drilled through the ball retaining sleeve 612, and project into the groove 602 machined on the valve stem sleeve 604. The ball retaining sleeve 612, at the point where the ball retaining holes 622 are drilled, is slightly thicker than one-half the diameter of the balls 600. The groove 602 in the valve stem sleeve 604 is slightly less than one-half the ball diameter.

The G weight 614 is held down over the balls 600, forcing them into the groove 602 in the valve stem sleeve 604, by the spring 618 mounted between the weight 614 and the shut-off valve cover 616. The valve stem 610 is threaded into the valve stem sleeve 604, and is provided with a knurled thumb screw 624 at the end. The valve stem sleeve 604 is spring loaded towards the open position by the spring 620 between the valve stem sleeve 604 and the ball retaining sleeve 612.

When the valve stem 610 is screwed into the valve stem sleeve 604, the valve disc 606 is seated on the seat 608. A force is applied to the balls 600 between the valve stem sleeve 604 and the ball retaining sleeve 612, locking the valve 446 in place. The balls 600 are held in position by the tapering inside diameter of the acceleration actuated weight 614 which fits over them.

The weight 614 is held in place by the precision spring 618. During launching, when the missile acceleration reaches between 5 and 12 g's, the G weight 614 moves against the pressure of the spring 618 to compress it to a position where the larger inside diameter of the weight 614 allows sufficient relief space for the balls 600 to move outward into the larger diameter machined on the inside of the weight 614. This movement of the six balls 600 releases the valve stem sleeve 604, from the groove 602, allowing the valve stem spring 620 and the high pressure gas on the bottom of the valve disc 606 to open the valve 446.

When the boost phase of the missile 100 is over, the G weight spring 618 moves the weight 614 back to its original position, forcing the balls 600 into a second groove 626 machined on the valve stem sleeve 604, thus locking the shut-off valve 446 in the open position. The valve 446 is sealed against leakage in the open position by an O ring 628 fitted into a groove 630 on the back face of a valve disc cap 632.

The valve seat 608 is an insert with a lip 634 machined on one surface 636 thereof. This lip 634 is positioned to mate with the valve disc 606. The valve seat 608 is secured in place with machine screws 638, and it is sealed with an O ring 640.

The valve disc 606 comprises a cup 642 and a cap 632, which are screwed together and locked by a lock washer 644. The cup 642 contains a resilient washer insert 647 which mates with the lip 634 machined into the valve seat 608 to seal against the high pressure gas applied to the bottom of the disc 606. The cap 632 of the valve disc 606 is fitted with the O ring 628 to seal against gas leakage past the valve stem 610 when the valve 446 is open.

A collar 646 is held between the cap 632 and cup 642 of the valve disc 606, and it is screwed and pinned to the valve stem 610. The shock absorbing washer or gasket 648 is sandwiched between the collar 646 and the cap 632. When the valve 446 is opened, the valve disc 606 moves up into the ball retaining sleeve 612, with the O ring 628 sealing against the valve disc cap 632 and the ball retaining sleeve 612. The hard rubber shock absorbing gasket 648 prevents excessive shock on opening of the valve from breaking the valve stem 610.

Pressure regulation by the flow control unit 444 is carried out by two nearly identical pressure regulation stages 650 and 652. Each pressure regulation stage 650 or 652 consists of a reference chamber 654 or 656 containing a spring 658 or 660, a regulating piston 662 or 664 containing ports 666 or 668, a ported piston sleeve 670 or 672 and a damping piston assembly 674 or 676, respectively. The outlet 678 of the nitrogen shutoff valve 446 is directly connected to an annulus 680 surrounding the first stage pressure regulating piston sleeve 670.

This sleeve 670 contains four ports 682, that are aligned with four triangular ports 666, cut through the regulating piston 662. The regulating piston 662 is held in the open position in the sleeve 670 by means of the 30 pound spring 658 mounted in the reference chamber 654. The second stage regulator outlet pressure is also applied to the first stage regulator reference chamber 654, by means of a bleed 651.

A damping piston 684 is inserted into the inside of the regulating piston 662. The damping piston 684 is held in a fixed position by means of a tie rod 686 securing it to the flow control unit body 560. When high pressure nitrogen from the shut-off valve 446 is applied to the first stage 650 of the pressure regulator 444, the regulating ports 666 in the piston 662 are open. Nitrogen passes through the ports 682 in the piston sleeve 670 and the ports 666 in the piston 662 to the inside of the piston 662. Nitrogen leakage past the damping piston 684 applies the first stage regulated pressure to the space between the top of the damping piston 684 and the closed end 688 of the regulating piston 662.

When the pressure applied to the inside end 688 of the regulating piston 662 exceeds that applied in the opposite direction by the reference chamber 654 and the spring 658, the piston 662 moves in such a manner as to close the ports 666 in the piston 662 and keep the total force applied to the top 688 of the piston 662 equal to that applied to the bottom of the piston 662. The spring 658 mounted in the reference chamber 654 acts over an area of approximately one-third square inch. Thus, the first stage regulator outlet pressure is approximately 100 psi higher than the outlet pressure of the flow control unit 444. The damping piston assemblies 674 and 676 are included in the unit 444 to damp out rapid pressure fluctuations on the bottom side of the pistons 662 and 664, and give a smoother outlet pressure.

The first outlet pressure of the first regulation stage 650 is applied directly to the inlet 690 of the second pressure regulating stage 652. The second pressure regulation stage 652 of the flow control unit 444 is identical to the first with the following exceptions noted below. The reference chamber 656 is vented to ram pressure and contains an accurate spring 660 that can be adjusted, and the walls of the piston 664 of the second pressure regulator stage 652 are thinner than in the first pressure regulation stage 650. As the spring 660 in the second pressure regulation state 652 acts over an area of slightly less than one-third square inch, the flow control unit outlet pressure is 28 psi above ram pressure.

The nitrogen flow control unit outlet 692 is connected to tubing 694 which leads to the relief valve 452 and bladders 438. A tubing fitting (not shown) also connects the flow control unit outlet 692 to the bladder vent fitting 552. This fitting is of the quick disconnect type which bleeds the bladders 438 to atmospheric pressure when installed. This vent 552 is fitted with a screen to prevent foreign particles from entering the nitrogen system. When the bladders 438 are removed, the male part of the quick disconnect fitting closes, preventing loss of nitrogen from the system. This vent fitting must always be removed before aerial flight of the missile 100.

For flight of the missile 100 at maximum Mach number, the air pressure of the diffuser inlet of the turbine fuel pump 440 is sufficient to prevent cavitation during missile operation below 19,000 feet altitude. For minimum flight Mach number, the diffuser air pressure is sufficient below 8,000 feet altitude. Since the nitrogen supply is rapidly used up in low altitude flight, the check valve 450 is placed in the nitrogen pressurizing system and is connected to the air turbine inlet air. When this pressure is greater than ram pressure plus 28 psi or whenever the nitrogen supply is insufficient to provide ram pressure plus 28 psi (as would occur in a long flight of the missile 100 at high altitude which terminated in a low altitude homing phase), the check valve 450 will open, permitting this pressure to inflate the fuel pressurization bladders 438.

This check valve 450 is of a hinge check type, and is installed transversely to the longitudinal axis of the missile 100 with the hinge forward to prevent the high boost acceleration of the missile 100 from opening the valve 450 with a corresponding loss of nitrogen.

The fuel pressurization system 228 also includes the safety nitrogen relief valve 452 for the protection of the fuel tank structure in case of failure of the nitrogen flow control unit 444, or temporary failure thereof produces excessive inflation pressures in the bladders 438. This valve 452 is located in line 694 between the flow control unit 444 and the inlet line 530 to bladders 438. As shown in FIG. 69, this safety relief poppet valve 452 is of the balanced piston poppet type and consists of a balanced piston 696 held against a circular seat 698 by a compression spring 700 and the scoop inlet pressure, in housing 702. The spring side of the piston 696 is referenced to the inlet pressure of the air-turbine fuel pump 440 by tubing 528 and the opposite side is open to nitrogen pressure through tubing 694 for the bladders 438. The differential pressure across the piston 696 is very nearly the differential pressure across the inner wall 454 of the fuel tank 138. A venting outlet 704 from the valve 452 is large enough to blow off the complete nitrogen supply in case of complete failure in the flow control unit 444.

In the event there is complete failure of the nitrogen flow control unit, that is, with all valves 446, 448, 650, and 652 wide open, the safety relief valve 452 would start to open when the differential pressure across it rose to 65 psi. The maximum area of the valve 452, is sufficient to limit the fuel tank wall differential pressure to 75 psi. The valve 452 discharges to the compartment 224 containing the nitrogen pressurization system 228, with complete blow-off of nitrogen being insufficient to damage the compartment 224 or the fairings around the fuel tank 138.

As shown in FIGS. 9A, 9B, and 13, the function of the fuel shut-off valve 428 is to seal the fuel system at the end of the outlet line 436 of the fuel tank 138 so that the tank may remain fully fueled for an indefinite period without leakage therefrom, and to permit normal operation of the fuel system after the missile 100 is launched. It is estimated that the valve 428 will be fully open 0.1 second after firing of the missile 100. It opens in a positive manner and is locked in the opened position during flight of the missile 100.

The fuel shut-off valve 428 is located in the fuel system between the outlet of the fuel tank 138 and the inlet 532 of the fuel pump 440. It is mounted to an inlet flange 706 of the fuel pump 440, by suitable means, such as bolts or the like. The fuel shut-off valve 428 comprises a body 708, a ball retaining sleeve 710, a G weight 712, a valve disc assembly 714, a valve opening spring guide 716, two springs 718 and 720, eight balls 722, a cover 724, and O rings 726 and 728.

The ball retaining sleeve 710 is drilled near one end 730 with eight holes 732 to contain the eight balls 722. At the point where these holes 732 are drilled, the thickness of the sleeve 710 is slightly more than half the diameter of the balls 722. The ball retaining sleeve 710 is held in the main body casting 708 by the valve cover 724 and is sealed against leakage by the O ring 728. A valve stem sleeve 734 fits inside the ball retaining sleeve 710 and has two grooves 736 and 738 machined in it. The depth of these grooves 736 and 738 is slightly less than one-half the ball diameter.

A valve opening spring guide 740 fits inside the valve stem sleeve 734 and is centered by a hub 742 machined on the bottom of the ball retaining sleeve 710. The valve opening spring 718 fits around this guide 740 and acts to force the valve disc 714 open. A valve stem 744 screws into the valve stem sleeve 734, and passes through a centering hole 746 in the ball retaining sleeve 710.

The G weight 712 fits around the outside of the ball retaining sleeve 710. Its inside diameter is so machined as to hold the balls 722 in the groove 736 machined in the valve stem sleeve 734. The G weight 712 is held in its normal position by a precision spring 720. This spring 720 forces the G weight 712 away from the main body casting 708.

The valve disc 714 is secured to the valve stem 744 by means of a threaded flanged collar 748, and a pin 750. The threaded collar 748 is held in place in the valve disc 714 by means of a shock absorbing soft rubber washer 752, a washer 754, and a snap ring 756. The sealing side 758 of the valve disc 714 is fitted with an annular resilient ring 760 so designed to seal tighter in its machined seat 762 in the body casting 708 with an increase inlet pressure of the fuel from tank 138.

Leakage around the valve stem 744, when the valve 428 is in the open position, is prevented by the O ring 726 on the downstream side of the valve disc 714 seating in a machined seat 764 in the ball retaining sleeve 710. Rotation between the valve stem sleeve 734, the ball retaining sleeve 710, and the G weight 712 is prevented by guide pins 765 and 767 riding in machined slots 769 and 773, respectively.

In the closed position of the valve 428, the locking balls 722 are held in the ball retaining sleeve holes 732 in a "closed position" locking groove 736 in the valve stem sleeve 734 by the G weight 712. Screwing the valve stem 744 into the valve stem sleeve 734 seals the valve disc 714 against its seat 762. Movement of the valve stem sleeve 734 relative to the ball retaining sleeve 710 is prevented by the balls 722. The valve stem 744 should be tightened finger tight to prevent leakage of the fuel and to insure proper loading on the balls 722.

When the missile 100 is launched, launching accelerations cause the G weight 712 to move against the pressure of the spring 720 toward the valve body 708. When the G weight 712 moves toward the valve body 708, a larger diameter on the inside of the G weight 712 is presented to balls 722. The valve opening spring 718 acting on the valve stem sleeve 734 causes the balls 722 to move radially outward into this increased diameter, and to move out of the groove 736 machined in the valve stem sleeve 734, unlocking said sleeve. The valve opening spring 718 causes the valve disc 714 to open and seat the O ring 726 against the ball retaining sleeve 710 and seals the stem 744 against leakage.

When the accelerations of the missile 100 decrease to near zero, the G weight spring 720 returns the G weight 712 to its normal position, forcing the balls 722 back into the second groove 738 machined in the valve stem sleeve 734, locking the valve disc 714 in the open position. The G weight 712 functions to unlock the valve 428 when launching accelerations of the missile 100 reach between 5 and 12 g's. The valve opening spring 718 is preselected to open the valve 428, with no fuel pressure applied to the valve inlet 434 against longitudinal accelerations up to 42.5g and to keep the valve 428 fully open against a 29g force. Once the valve 428 is locked open, no acceleration forces applied to the missile 100 can cause it to close.

The fundamental difference from the nitrogen shut-off valve 446 and the fuel shut-off valve 428 exists in that packaging requirements for the nitrogen G valve 446 have forced the opening spring 620 of the valve 446 to work against the inertia forces acting on the valve disc 606 and stem 610. In addition, no fuel pressure is counted upon to aid in opening the valve 446.

The triggering weight 712 for valve 428 weighs approximately 1.6 pounds, with the design opening point being from 5 to 12 g's. As indicated, the valve opening spring 718 is capable of opening the valve 428 against 42.5g and will hold the valve 428 fully open against 29g until the acceleration of the missile 100 had decreased sufficiently for the G weight 712 to return to its original position and key the valve open. The O ring 726 fitted into the groove 766 on the back side of the valve disc 714 provides positive sealing against leakage when the valve is in the open position.

After separation of the missile 100 from the booster 500, no maneuvering of the missile can cause the valve 428 to close. However, during ground testing of the missile 100, the valve 428 may be activated manually and reset quickly innumerable times.

This fuel shut-off valve 428 also incorporates a torque limiting thumb screw 768 for adjustment of the valve disc 714. This screw 768 will provide 12 to 14 pounds end thrust of the valve disc 714 against its seat 762.

As best seen in FIGS. 9A and 12, the centrifugal fuel pump 440, which is driven by a turbine 770, is the primary source of fuel pressure required for regulation and injection of the fuel in the combustor 108. This pump 440 provides operating heads for the various fuel components. The air turbine 770 is driven by air received by the scoop 772 in the turbine air intake strut 774 in the diffuser 270.

The fuel pump assembly 440 consists of a cast aluminum housing 778, the air turbine 770 with forty blades 784, an air turbine nozzle plate 786 containing eighteen nozzles 456, a centrifugal fuel pump 776 with four vanes 788, and a bearing and seal assembly 790. The air turbine 770 and the centrifugal pump 776 are mounted at opposite ends of a common shaft 780 and are sealed from each other by a high speed carbon faced mechanical seal 782.

The air turbine 770 consists of a single stage impulse wheel 771 driven by air taken from the center of the diffuser 270 by the air intake scoop 772 in the air intake strut 774. The air enters the wheel 771 from the eighteen converging-diverging supersonic nozzles 456, which have a designed exit Mach number of 1.58 and a throat diameter of 0.250 inch. These nozzles 456 are mounted in the air turbine nozzle plate 786, which is bolted to the pump body 778. The entire pump unit 778 is mounted in the missile 100 with the nozzle plate 786 fitting into the air intake scoop 772 in the strut 774. The air is discharged from the turbine wheel 771 into a pump compartment 792, and led overboard by means of large louvers (not shown) installed in the missile skin 286.

The turbine 770 is mounted, as previously mentioned, on a tapered shaft 780 and is held in place by a special torque washer 794 fitted to flats (not shown) on the shaft 780 and identations (not shown) on the turbine wheel 771. The torque washer 794 is held in place by means of two right-hand thread lock nuts 796.

The impeller 798 for the pump 440 is of the four vaned, single stage, single entry shrouded type and is mounted on the end of the pump shaft 780 opposite the air turbine 770, as previously indicated. The impeller 798 discharges into an involute shaped cavity case 800 in the main body housing 778. It is keyed to the pump shaft 780, and held on the shaft by means of a streamlined lock nut 802. The impeller 798 is pressure balanced to reduce axial thrust loads on the bearings 804 and 806 to nearly zero. The bearing and seal assembly 790 is mounted in the body casting 778 between the turbine wheel 770 and the pump impeller 798.

The two ball bearings 804 and 806 are mounted on the shaft 780, with one being located at each end of the bearing and seal assembly 790. Each of these bearings 804 or 806 is designed to take radial thrust in all directions and axial thrust in one direction. These bearings 804 and 806 are mounted back to back so that axial thrust in either direction is taken up. The bearings 804 and 806 are held in place by steel snap rings 808 and 810, respectively.

The carbon seal assembly 782 is mounted between the bearings 804 and 806. The seal 782 consists of a flanged sleeve 812 rotating with the shaft 780 and an annular spring loaded carbon seal 814 which is mounted in the bearing and seal assembly 790. It rides against the flange 816 on the rotating sleeve 812. This seal 814 prevents leakage of the fuel past a pressure balancing seal 818 and through the impeller bearing 804 down the shaft 780 to the turbine bearing 806 and the turbine exhaust air stream. Leakage past the impeller bearing 804 is vented through line 820 back to the inlet side 532 of the pump 440.

The impeller bearing 804 is cooled and lubricated by the fuel leaking past it to the carbon seal 814. The turbine bearing 806 is packed in a high temperature grease for lubrication. The entire bearing and seal assembly 790, as indicated, is mounted in the body casting 778 and is sealed against fuel leakage between the assembly housing 822 and the body casting 778 by O rings 824 and 826. The bearing and seal assembly 790 is held in place by a large spanner nut 828 which fits over the bearing and seal housing 822 and is screwed and lock-wired to the central portion 830 of the body casting 778.

The discharge pressure of the pump 440 is between 300 psi and 400 psi at sea level depending on flight velocity of the missile 100. The speed of the pump 440 is between 20,000 and 30,000 rpm at sea level under expected operating conditions. The maximum design rotational speed of the pump 440 is 40,000 rpm. The turbine wheel is spin tested under load to 50,000 rpm during check out of the pump 440.

The fuel strainer 458 is also provided in the fuel control system 230 to prevent contaminating particles from entering both the pilot and main fuel systems.

As shown in FIG. 10, this fuel strainer 458 includes a housing 832 which is integral with the main fuel regulator 380. One end 834 of the housing 832 is externally threaded to receive a threaded ring 836 for securing a flange 838 of an accordian type flexible connector 840 thereto. This connector 840 connects the outlet 842 from the fuel pump 440 to the fuel strainer 458. The other end 844 of the housing 832 is internally threaded to receive an end cap 846. This end cap 846 is utilized to close end 844 of the housing, after a strainer element 848, which consists of a hollow sintered bronze cylinder, having a length of 6.875 inches, with outside and inside diameter of 2.438 inches and 2.094 inches, respectively, is inserted in the housing 844. A coiled spring 850, located between a centering element 852 and the end cap 846, is utilized for centering the strainer element 848 in the housing 832.

The fuel from the fuel pump 440 enters the fuel strainer 458 through end 834 of the housing and is passed through the strainer element 848. The fuel then filters outward through the strainer element 848 to an annulus 854 surrounding the element 848. This annulus 854 feeds into the entrance passages 856 and 858 of the main and pilot fuel regulators 380 and 340, respectively. The strainer element 848 is capable of trapping all particles with a dimension greater than 0.001 inch. O rings 860, 862, and 864, are utilized to seal the fuel strainer 458 from leakage of fuel therefrom.

The pilot fuel regulator 340 is designed to provide the required fuel flows to the pilot section 320 of the combustor 108 to insure a stable pilot frame which is not affected by the velocity of the missile 100. The pilot section 320 of the combustor 108 uses about 20 percent of the air flow passing through the combustor, and it operates at a nearly constant equivalence ratio of 1.4 which insures stable operation of the combustion system at all times. The fuel flow for the main combustor is influenced by a velocity control mechanism. It may exceed the limits required for stable operation, with the result that the ramjet flame may be extinguished without the stabilizing influence of the pilot section 320. This is especially true at high altitudes.

The air-fuel ratios are rich to insure stable operation of the combustion system. However, it has been determined that a leaner mixture is most suitable for ignition. During the starting period, use of a lean mixture helps to eliminate or reduce violent explosions. In order to obtain a leaner mixture in the pilot section 320, a squib activated unit 866 allows the pilot regulator 340 to deliver an air-fuel ratio of 0.0663 until ignition of the fuel takes place shortly after separation of the missile 100 from the booster rocket 500. One to 3 seconds after ignition of the fuel, the squib unit 866 operates, raising the air-fuel ratio to the normal rich value.

In order to obtain the required air-fuel ratios, the pilot regulator 340 must deliver fuel flows which vary from 4.5 lb/sec to less than 0.3 lb/sec depending on altitude and speed. Air-fuel ratios must be held to within .+-.10 percent of the required values from sea level to 30,000 feet, and within .+-.5 percent from 30,000 to 60,000 feet.

The pilot regulator 340, which delivers the required fuel flow, derives its intelligence from two pressure pick-up points on the missile 100. Referring now to FIGS. 7, 9A, 9B, and 11, of the drawings, the ram pressure pick-up probe 128 is located at the forward end of the inner body 120, with the ram pressure being piped by aluminum tubes 868, 870, and 872, to the pilot and main regulators 340 and 380, and the nitrogen flow control unit 444. Atmospheric pressure is picked up by four body static taps 130, located on the missile skin on the forward body assembly 102. The pressures picked up from these four ports 130, are manifolded together and piped through aluminum tubes to the main and pilot regulators 340 and 380. Air flow through the ramjet engine is approximately proportional to ram pressure over the designed Mach number range. The fuel flow delivered in proportion to ram pressure will, therefore, maintain a desired and nearly constant air-fuel ratio.

This pilot regulator 340 is comprised of an altitude controlled variable area orifice and a pneumatically referenced pressure regulating linkage which controls the fuel pressure drop across the variable area orifice. Since the fuel is considered incompressible, the pilot regulator 340 thereby becomes independent of the outlet and inlet pressures provided the inlet pressure is sufficiently greater than the outlet pressure for the throttling type pressure regulating linkage to operate.

The variable area orifice consists of a contoured valve 874 whose travel is controlled by a servo type positioning mechanism 876. The fuel inlet pressure of the pilot regulator 340 is utilized as the high pressure source for the servo positioning mechanism 876, while atmospheric pressure is used as the low pressure source for the servo positioning mechanism.

A pilot valve 878 of this mechanism 876 is connected to an evacuated altitude bellows assembly 880 which senses altitude static pressure. Motion of the valve 878 is then a function of altitude. There is a negligible error due to the influence of fuel inlet pressure on the servo positioning mechanism 876. The required contouring of the valve 874 is determined by finding the required area from the characteristics of the pressure regulating linkage and the required fuel flow variation with altitude, and coordinating this area with the altitude-travel relationship of the evacuated bellows assembly 880, as indicated in Equation (3) below: ##EQU1## where A = area in square feet;

F.F. = required fuel flow pounds per second;

C = coefficient of discharge of the valve;

g = acceleration due to gravity in feet per second per second;

d = fuel density in pounds per cubic feet; and

.DELTA. p = controlled differential pressure across the valve in pounds per square feet.

The pressure regulating linkage is comprised of a poppet valve 882 which is controlled by two opposed equal area fuel and air pistons 884 and 886, respectively. A reference differential pressure is imposed across the air piston 886 by the pneumatic control circuit, and the resulting force is opposed by a force generated on the fuel piston 884 by the differential pressure across the variable area contoured valve 874. Any unbalance causes motion of the poppet valve 882, with a corresponding change in the amount of throttling of the fuel flow.

The change in fuel flow alters the differential pressure across the variable area contoured valve 874, and thereby alters the force of the fuel piston 884. The poppet valve 882 will move until equilibrium of forces is established. Such motion will occur whenever the differential pressure of the air piston 886 changes or whenever the inlet or outlet pressures of the pilot regulator 340 change. The differential pressure of the variable area altitude valve 874 is thus controlled and is equal to the differential pressure of the air piston 886.

At any constant altitude of the missile 100, the fuel flow is arranged to vary with ram pressure. Since the area of the metering orifice 888 in the variable area altitude valve 874 is fixed (constant altitude), the variation must be provided for by the differential pressure of the air piston 886. The constant of proportionality may be a function of altitude as this can be accounted for in the contour of the variable area altitude valve 874.

Since, at constant altitude, the fuel flow is proportional to the square root of the pressure differential across the variable area altitude valve 874 and the pressure differential of the air piston 886 equals the pressure differential of the variable area altitude valve 874, the pneumatic control circuit is required to produce a differential air pressure whose square root varies with ram pressure.

This is accomplished by a pair of restrictions 890 and 892 in series in a pneumatic control circuit 894 which pass air from ram pressure to body static pressure. The square root of the pressure drop across the downstream restriction 892 will very closely approximate the desired proportion over the range of applicable Mach numbers of operation of the missile 100.

Since ramjet ignition must occur quickly at separation, the time delay unit 866 is provided as a modification to the pneumatic control circuit 894. During the boost phase of the missile 100, and the separation thereof from the booster 500, an additional passage 896 is opened in parallel with the downstream restriction 892 of the control circuit 894 in order to lower the air piston differential pressure below its normal value. This conserves fuel during the boost phase after the fuel shut-off valve 428 has opened and provides a leaner mixture to the pilot section 320 of the combustor 108 for more favorable starting of the ramjet engine. At separation of the missile 100 from the booster 500, two instantaneous squibs in the time delay unit 866 explode to close a valve which shuts off this parallel restriction 896, with the result that the differential pressure of the regulator air piston 886 then quickly returns to its normal condition in approximately 0.8 seconds.

For purposes of clarity, the pilot regulator 340, FIG. 11, has been illustrated semi-schematically and various minor details have been deleted. The pilot regulator 340 has a body casting 898. Various channels are provided in the casting walls for transferring the fuel and air pressures used in the regulator 340. In actual practice, bosses are provided on the castings to permit pressure pick-offs to be used for testing and calibrating the regulator. All moving parts within the regulator subject to wear or corrosion are made of stainless steel. Two types of gaskets are used in the pilot regulator 340 to prevent air and fuel leakage, namely, flat type gaskets and O ring type gaskets. O rings are used in various sizes through the pilot regulator 340.

In practice, a poppet valve sleeve (not shown) is fitted into the fuel inlet side of the pilot regulator body casting 898. It is made of a stainless steel piece and it is sealed at each end by O rings. Four ports cut around the sleeve permit fuel flow through the sleeve. The top edge 900 of the poppet valve 882, which is wedged shaped to reduce Bernoulli effects, meters the fuel flowing through the sleeve ports. The bottom end of the valve is blanked off except for two apertures 902 and 904. One aperture 902 is used for attaching the valve 882 to the valve rod 906; while the second aperture 904 is used to equalize the pressures above and below the blanked end of the valve. Slots are cut in the bottom of the valve to reduce weight thereof and G forces in flight. These slots are cut in the valve 882 below the level of lower edges of the fuel ports cut in the valve sleeve when the poppet valve 882 is closed. A sleeve cover holds the sleeve in place in the pilot regulator body by means of lock wired screws, and it is sealed against fuel leakage by an O ring.

A baffle, at 908, is provided in the regulator body directly above the poppet valve sleeve. The baffle 908 reduces turbulence and local high velocity jets of fuel from the metering edge 900 of the poppet valve 882 and delivers an even pressure to the bottom of the fuel piston 884.

The clearance between the disc and the regulator casting 898 provides means for applying partially metered fuel pressure (altitude valve inlet pressure) to the bottom of the fuel piston 884. The center of the baffle disc is drilled for the fuel piston rod. Fuel flows from the poppet valve 882 into the open end of the baffle 908, then radially through the perforated walls to a fuel collecting annulus. Clearance between the baffle disc and the pilot regulator body casting 898 allows fuel as near static pressure as is possible, to be applied to the bottom 910 of the fuel piston 884.

A spacer fits into the regulator body above the baffle. The fuel piston sleeve contains a shoulder at the top. An O ring prevents fuel leakage between the sleeve, spacer, and regulator body 898. The fuel piston end of the rod 906 is drilled and tapped and the piston is bolted to the rod. The other end of the rod 906 is threaded for the poppet valve 882. The poppet valve 882 is attached by first placing a spring bushing 912 on the rod 906, over which is placed a poppet valve spring 914. The poppet valve 882, metering edge 900 toward the fuel piston 884, is then placed on the shaft bearing against the spring 914. Double washers 916 of the ball and socket type are then placed on the rod 906 and bear against the poppet valve 882. The assembly is held together with a lock nut 918. The ball and socket washers 916 and the spring 914 prevent any misalignment of the rod 906 from binding the valve 882.

The air piston housing, shown here as being cast integral with the regulator casting 898, is in actual practice a cast aluminum housing. The top end of the housing is open. The lower end 920 of the housing is closed except for a hole 922 in the center. A force pin 924 fits in a bushing (not shown) pressed in this hole 922 with a minimum clearance to permit frictionless movement but to prevent excessive fuel leakage. The low side of the air piston 886 communicates with the body static ports 130 by means of passage 926 and tubing 928.

A air piston sleeve (not shown) fits into the air piston body, with an "O" ring seal being utilized to prevent air leakage between the body and a shoulder of the sleeve.

The air piston 886 is, hand lapped to the sleeve. The air piston 886 bears against the force pin 924 which transfers the motion of the air piston 886 to the fuel piston 884 and poppet valve 882.

The air piston has a cover (not shown). Four bolts secure the cover and air piston housing to the regulator body 898. Flat gaskets provided between the regulator body 898, the air piston housing, and air piston housing cover seal the regulator 340 against fuel and air leakage.

The regulator 340 is provided with two passages 930 and 932 through which air flows from the ram pressure pick-up 128 to the atmospheric pressure outlets 130. These passages 930 and 932 are provided at one end with the bleeds 890 and 892 respectively. The inlet passage 930, connected to the ram pressure pick-up 128, contain bleed 890. The outlet passage 932, connected to the body static pressure ports 130, contains bleed 892. The pressure generated in the passage between these bleeds 890 and 892 is applied to the top (high pressure) side of the air piston 886. The downstream bleed 892 is tailored to give proper air heads. In addition to the inlet and outlet passages 930 and 932, the air piston communicates with the time delay unit 866 by means of passage 896 and tubing 934.

It is to be noted that the air piston 886, fuel piston force pin 924 and fuel poppet valve 882 are termed the fuel metering linkage. The fuel downstream of this linkage is termed partially metered fuel.

A sleeve 936 is provided in the variable area contoured valve 874. This sleeve 936 is mounted into the regulator body casting 898 on the downstream side of the fuel metering linkage. Four ports 938 are cut through the sleeve 936 to permit passage of fuel through the sleeve. Each end of the sleeve 936 is sealed against fuel leakage between the sleeve and the body casting by 0 rings (not shown).

The variable area altitude valve 874 (variable area orifice valve) is a piston type valve. The bottom 940 of the valve 874 is open. The top 942 of the valve 874 is completely closed, and is tapped (shown integrally in FIG. 11) to receive a servo follower rod 944. Six ports (only one 888 of which is shown) are cut in the valve 874 to permit fuel flow through the valve.

The ports, such as port 888, are designed to give proper fuel flows at each altitude with air piston differential pressure across the valve 874. The ports are 0.620 inch long, with the top of the ports 0.825 inch from the top 942 of the valve 874. When the valve 874 is fully open, the maximum port area is aligned with the sleeve ports 938, and maximum fuel flow results. As the valve 874 is closed, it is moved toward the regulator outlet 946, and the port area is decreased. The position of the valve 874 is a function of altitude. This will be discussed more in detail subsequently. To prevent the valve 874 from turning in its sleeve 936, a pin (not shown), secured to the sleeve 936, rides in a slot (not shown) machined parallel to the axis of the valve 874.

The variable area altitude valve 874 is provided with a coil spring 948 carried by a spider 950. This spring 948 applies a force to the altitude valve 874 that tends to open the valve. This force is overcome by the action of the altitude valve servo mechanism 876 described below.

A fuel outlet elbow 952, FIG. 8, is bolted to the pilot regulator body 898 below the altitude valve assembly 874. A flat gasket is utilized to seal the body casting 898 and the outlet elbow 952 against fuel leakage.

A regulator outlet pressure pick-up 954 is mounted in the regulator outlet 946. An end 956 of the pressure pick-up 954, which projects into the fuel stream, is closed. Two apertures 958 drilled in the side of the tube are used to pick up the pressure. The other end 960 of the pressure pickup 954, which is secured to the regulator outlet 946, is sealed with an 0 ring (not shown). It is connected to the low pressure side of the fuel piston 910 by passage 962 provided in the body casting 898.

The altitude bellows 880 is of the sealed, evacuated, spring loaded type. The free length of the bellows 880 at sea level is approximately 1 1/2 inches. One end 964 of the bellows 880 is fixed, while the other end 966 is free to move. The bellows 88, as mounted, has atmospheric pressure applied to the space surrounding it so that with a decrease in atmospheric pressure, it will increase in length.

The free end 966 of the bellows 880 is connected by a thin rod 968 to an altitude valve servo piston 970, and moves about 0.018 inch per inch mercury change of barometric pressure. The bellows 880 will increase approximately one-half inch in length with a pressure change surrounding it corresponding to an altitude change from sea level to 60,000 feet. The fixed end 964 of the bellows 880 is closed by means of a cap 972 that is threaded for adjustment purposes. The cap 972 is locked with a nut 974 after adjustment.

The servo mechanism pilot valve 876 is designed to position the variable area altitude valve 874 with respect to the free end of the altitude bellows 880 and to make the motion of the contoured altitude valve 874 (variable area valve) correspond exactly to the motion of the altitude bellows. The high pressure of the fuel at the inlet 976 of the pilot regulator 340 is the actuating medium for this pilot valve. Under equilibrium conditions, high pressure fuel fills an annulus 978 between the servo piston 970 and the servo follower 980 by flow through a servo follower inlet annulus 982. The servo piston 970 is pressure balanced by means of a hole 984 through the pilot valve to eliminate forces on the piston ends.

As altitude of the missile 100 is increased, the altitude bellows 880 expands. Since the rod 968 is linked to the servo piston 970, it moves the servo pilot valve 878 toward the altitude valve 874. For any given altitude, the position of the altitude bellows 880 and servo piston 970 is fixed. this motion of the servo piston 970 uncovers ports 986 in the servo follower 980, venting high pressure fuel through the ports 986 to the top 942 of the altitude valve 874. Venting of the high pressure fuel to the top 942 of the altitude valve 874 causes this valve to move downward (toward the closed position).

As the valve 874 moves, it carries the servo follower 980 with it due to the rod 944 between them. As the servo follower 980 moves, the high pressure ports 986, carrying fuel to the top 942 of the altitude valve 874, move under the lands 988 of the servo piston 970, and the flow of fuel is shut off, stopping the motion of the altitude valve 874. As the missile 100 decreases in altitude, the altitude bellows 880 contracts, and through the rod linkage 968, mves the servo piston 970 away from the altitude valve 874. This motion of the servo piston 970 uncovers the ports 986 in the servo follower 980 and vents pressure from these ports to an overboard drain 990 by means of a tube 992. As the pressure is released from the top 942 of the altitude valve 874, the force of the altitude valve spring 948 and the regulator outlet pressure move the altitude valve 874 toward the open position, causing the servo follower 980 to move so as to cover the ports 986 in the servo follower with the pilot valve lands 988.

In actual practice, the servo mechanism 876 is carried by a servo sleeve (not shown) pressed into the servo housing, which is, in turn, bolted to the main body casting 898. Passages 944 and 946 in the regulator 340 permit flow of fuel pressures to and from the servo mechanism 876. The servo follower 980 is a hollow spool type valve. The servo piston 970 is a spool with two lands 988, one at each end, which are hand lapped to the inside diameter of the servo follower 980.

as illustrated in FIg. 70, the squib activated time delay unit 866 comprises a body 988 containing a spring loaded valve 1002 held in the open position by a fine wire 1004. The inlet 1006 to the unit 866 contains a bleed (not shown) vented to atmosphere which parallels the outlet bleed 892 in the air piston circuit when the spring loaded valve 1002 is in the open position. The use of the aforementioned bleed to parallel the outlet bleed 892 causes a reduction in air piston differential pressures, and reduces the fuel flow to the valve necessary to give a 0.0663 air-fuel ratio for starting. Shortly after the missile booster separation, a time delay relay ignites an instantaneous squib 1006 which severs the fine wire 1004 holding the time delay unit 866 open, permitting the valve 1002 to close and take the bleed at inlet 1006 out of the circuit. This returns the pilot regulator 340 to normal operation.

The operation of the pilot regulator 340 will not be explained, starting with an explanation of the missile 100 on the launcher 1000. While the missile 100 is on the launcher 1000, no air pressures or fuel pressures are imposed on the pilot regulator 340, and consequently no fuel flows. During boost of the missile 100 by the booster 500, ram pressure is built up, creating air heads on the air piston 886. These air heads are lower than normal due to the bleed in the squib activated time delay unit 866 being parallel with the air piston air circuit. During boost, therefore, fuel flow has started, and fuel has filled the lines to the pilot regulator 340, the main regulator 380, and the lines to the fuel injector systems 320 and 378. By the time of separation of the missile 100 and the booster 500, fuel has started to flow from the pilot injectors 330, and the pilot igniter 1008 is ignited.

After separaton of the missile 100 from the booster 500 and ignition of the pilot igniter 1008, the squib activated time delay unit 866 is fired, and the bleed is taken out of the air circuit so that air heads and fuel flow go to normal design values. the altitude valve 874 has been moved to the proper position by the action of the altitude bellows 880 and the altitude servo mechanism 876. The fuel flows, as necessary, to give an air-fuel ratio of 0.0934 (up to 40,000 feet).

Let it be assumed that at an altitude of 5,000 feet, the missile 100 flies a level course, and it has a velocity of 1,900 ft/sec. The pilot air-fuel ratio will be at 0.0934. The main regulator 380 will be delivering enough fuel to accelerate the missile 100. As the missile 100 accelerates to 2,000 ft/sec, the ram air pressure increases to increase the pressure on the high side of the air piston 886. The pressure on the low side will be constant at body static pressure. The higher differential pressure across the air piston 886 moves the fuel metering linkage in such a manner as to impose the same differential pressure across the variable area altitude valve 874. As fuel flow through this valve 874 is proportional to the square root of the differential pressure (air piston differential pressure) across it, fuel flow through the pilot regulator 340 increases, as necessary to maintain a constant air-fuel ratio.

Let it now be assumed that the missile 100 is flying at 45,000 feet, with a speed 2,000 ft/sec. As the missile 100 climbs, ram pressure, static pressure, and fuel inlet pressure will decrease, the air iston 886 differential pressure will also decrease. At the same time, the pressure at the fuel inlet 976 will decrease, and the fuel metering linkage will move in such a manner as to impose air piston differential pressure across the variable area altitude valve 874. As the static pressure decreases, the altitude bellows 880 and servo mechanism 876 will cause the altitude valve 874 to move toward the closed position, reducing the port area available for fuel flow. The altitude valve ports 888 have been so designed to give a fuel flow corresponding to an air-fuel ratio of 0.0868. Since fuel and air flow through the missile 100 are both proportional to the square root of air piston differential pressure, this air-fuel ratio will be maintained at all velocities at this altitude.

In order to insure that adequate airheads are provided on the pilot regulator 340 of sufficient magnitude to overcome the acceleration forces during boost of the missile 100 and to insure fuel flow from the pilot injector at separation of the missile from the booster 500, regulated nitrogen pressure (pressure plus 28 psi) can be bled into the pilot regulator ram line. If this is done, a check valve should be added in the ram line to the pilot regulator 340 to prevent loss of nitrogen through this line, and a bleed should be installed in the regulated nitrogen line to the pilot regulator ram connection to adjust the nitrogen pressure applied to the high pressure side of the air piston. This bleed should be chosen so as to give the normal operating air heads on the air piston (normal rich pilot fuel flow) at separation. The pressures applied by the regulated nitrogen and the rate of pressure rise is such that, at separation of the missile 100 from the booster 500, the tube volume between the pilot regulator 340 and the pilot injector is filled, and fuel flows from the pilot injector.

At separation of the missile 100 from the booster 500, a squib activated valve can be utilized to shut off the regulated nitrogen to the pilot regulator 340. At this time, ram air opens the ram line check valve mentioned above, and the pilot regulator 340 now operates normally. It will be noted that at this time, the lean start time delay unit is in the circuit, and the pilot regulator 340 immediately goes to lean start operation. At approximately 2.2 seconds after separation of the missile from the booster, the lean squib unit operates, and returns the pilot regulator 340 to normal, rich stable burning operation.

The main fuel regulator 380, as shown in FIGS. 9A, 9B and 10, controls the delivery of the fuel to the main section 376 of the combustor 108 of the ramjet engine at various flight conditions of velocity and altitude, and operates in such a manner as to satisfy the engine requirements using MIL-F-5616 kerosene or the equivalent. This regulator 380 controls fuel flows within the desired limits from sea level to 60,000 feet altitude so as to produce a missile flight air speed of 2,000 ft/sec .+-. 5 percent by raising or lowering the fuel flow to the main section 376 of the combustor 108, thus increasing or decreasing thrust as required.

The fuel flow must vary with ram pressure at a constant altitude from a Mach number equal to 1.725 to the lower velocity control limit. The regulator 380 is programmed with altitude to provide an equivalence ratio of approximately 0.8 to the main section 376 of the combustor 108. The acceleration of the missile 100 will be reduced until zero acceleration or constant velocity flight of the missile results. Extreme climb, dive, or maneuver, where the full range of fuel flows may not be sufficient to produce a zero acceleration condition of the missile 100, may cause the missile 100 to depart from the controlled velocity region until such flight conditions have been attenuated. Such conditions, however, will be of very short duration.

Velocity control of the missile 100 is accomplished, as previously mentioned, by sensing the stagnation temperature, ram pressure, and static pressure. Equation (4) below illustrates how these quantities determine the velocity of the missile 100: ##EQU2## where P.sub.r /P.sub.a = ratio of ram to static or ambient pressure;

V = velocity in feet per second;

T.sub.s = stagnation temperature in degrees Rankin; and

K = ratio of specific heats at constant pressure and volume (assumed as 1.4 for air).

Air at ram pressure is fed to a circuit consisting of a fixed inlet bleed 1010 and a variable area outlet bleed 1012 and it is discharged at static pressure. The variable area bleed 1012 is powered by a bellows 1014 so that its motion is proportional to the differential pressure across the bellows. This differential pressure is the output of a temperature pick-up unit 1016 consisting of a thick, high response bimetallic element 1018 projecting into the diffuser 270 of the missile 100 which controls the pressure drop across a light poppet valve 1020 in series with a small bleed 1022. The output of this device 1016 is essentially proportional to stagnation temperature and independent of altitude and the pressure of the diffuser 270 as indicated by Equation (5) below:

P = T.sub.s /90 Eq. (5)

where

P = differential pressure output in pounds per square inch; and

T.sub.s = stagnation temperature .degree.F.

The variable area bleed 1012 is contoured in such a manner that the pressure ratio across it, which is determined by stagnation temperature and the ratio of ram to atmospheric pressure, becomes approximately a simple linear function of velocity, as indicated by Equation (6) below:

P.sub.x /P.sub.a = 0.00605V - 8.3 Eq. (6)

where

P.sub.x /P.sub.a = pressure ratio across the downstream (contoured) bleed; and

V = velocity in feet per second.

This pressure ratio is applied to an evacuated double Mach sensing bellows assembly 1024 which has a fixed end 1026 and a free end 1028 which is connected to a balanced air piston shorting valve 1030. The behavior of the Mach sensing bellows 1024 is such that at a pressure ratio equal to the area ratio of the two bellows 1033 and 1035, the free end 1028 is always in the same position. A greater pressure ratio causes motion in one direction and a lesser pressure ratio causes motion in the opposite direction. The area ratio of the Mach sensing bellows 1024 is equal to the pressure ratio existing across the contoured bleed 1012 at the lower design limit velocity, with the valve 1030 being set on the verge of opening at this condition. At all lower velocities, the valve 1030 remains closed, but as the lower design velocity is exceeded the valve 1030 begins to open. The valve 1030 is fully open by the time the upper design velocity is reached.

A small piston 1032 is also connected to the free end of the Mach sensing bellows 1024. It is fed by the high side pressure of the air piston 1034 and a separate double restriction air circuit 1036. Its purpose is to help equalize the sensitivity of control with changes in altitude of the missile 100 by preventing excessive travel of the valve 1030, with a given velocity overshoot at low altitude, thereby allowing high altitude sensitivity to be increased while holding a given maximum sea level sensitivity. The Mach sensing bellows 1024 is adjustable such that the presently specified 2,100 ft/sec control velocity is achieved within .+-. 100 ft/sec limits.

The ram (free stream pitot total) pressure probe 128 supplies one of the necessary intelligence measurements for the fuel regulators 340 and 380, and it is located, as previously mentioned, at the stagnation point at the forward tip of the inner body 120. The aluminum tube line 868 leads back through one of the center body support struts 324 to the fuel regulation compartment 226.

Air flows from the control circuitry of the fuel regulators 340 and 380 are discharged to the aluminum tubing line lines 928, 1038, 1040, and 1042 which run to the four body static vents 130 in the external surface of the missile 100 where the forward body section 102 and the center body section 106 come together. Due to the air flow in the body static lines 928, 1038, 1040 and 1042 and the external flow pattern of the missile 100, the static pressure in these lines in the vicinity of the regulators 340 and 380 is not an accurate intelligence of static pressure.

A velocity control mechanism 1044 which requires an accurate static pressure measurement obtains this measurement from a static probe 126 extending in front of any probable flight normalshock position. There is no flow in a line 1046 which connects the probe 126 to the velocity control mechanism 1044, but a pressure relief valve 1048 is provided to relieve the high pressure generated during the boost phase of aerial flight of the missile 100 when a shock wave is pushed out in front of the missile to such a location that the pick-up apertures in the static pressure probe 126 are in a region of higher than static pressure. This pressure relief valve 1048 is of a ball-check relief type, and requires a differential opening pressure of 4.5 to 7.5 psi. It is located in the fuel system compartment 226, and is, therefore, referenced to the pressure in this compartment.

In FIG. 10, the main fuel regulator 380 comprises a main body casting 1050 in or on which are mounted the following assemblies: the inlet fuel strainer 458, a sensitivity equalizer 1052, the Mach sensing bellows assembly 1024, the air piston shorting valve 1030, the air piston 1034, a fuel piston 1054, a fuel flow straightener 1056, a fuel poppet valve 1058, the previously mentioned velocity sensing bellows 1014, the variable area bleed 1012, an altitude sensing bellows 1060, a variable area orifice servo 1062, and a variable area orifice valve (altitude valve) 1064.

The Mach sensing bellows 1024 is an evacuated double bellows assembly consisting of the two bellows 1033 and 1035. One end 1026 of the Mach sensing bellows 1024 is fixed to a fitting 1068 and the other end 1028 is free. The free end 1028 of the Mach sensing bellows 1024 is connected to the air piston shorting valve 1030 by a thin force pin 1070. The output pressure from the velocity sensing unit 1072 is applied to the inner bellows 1035 through a passage 1074 provided in fitting 1068, and atmospheric pressure is applied to the outer bellows 1033 by means of tube 1046.

The area ratio between the inner and outer bellows 1035 and 1033, respectively, at the free end of the Mach sensing bellows 1024 is equal to the ratio of ram pressure to atmospheric pressure at a Mach number equal to 1.6; that is, the area of the outer bellows 1033 is 3.8 times the area of the inner bellows 1035. At pressure ratios lower than that corresponding to approximately a Mach number of 1.4, the free end 1028 of the Mach sensing bellows 1024 rests on mechanical stops (not shown) built into the Mach bellows mounting.

At pressure ratios above that corresponding to a Mach number equal to 1.4, the Mach sensing bellows 1024 increases its length. At a Mach number of 1.6, the length of the Mach sensing bellows 1024 is equal to that of an unevacuated bellows. Increasing pressure ratios causes the Mach sensing bellows 1024 to further increase its length. As the air piston shorting valve 1030 is mechanically linked to the free end 1028 of the Mach sensing bellows 1024, its motion corresponds directly to that of the Mach sensing bellows.

The air piston shorting valve 1030 is comprised of a spool 1076 located in a valve sleeve 1078. It is positioned by a spring 1080 located at one end of the valve sleeve housing 1082 and the force pin 1070 from the free end of the Mach sensing bellows 1024 at the other end. The valve spool 1076 is pressure balanced by means of a longitudinal passage 1084 through its center which vents the outlet pressure of the velocity sensing assembly 1072 from passage 1086 to both ends of the spool 1076, thereby eliminating any pressure unbalance on the spool. The valve sleeve 1078 is mounted in the base of the air piston assembly cover 1088.

The sleeve 1078 contains two annular ports 1090 and 1092. Port 1090 connects directly to the high pressure side of the air piston 1034 through passage 1094 while the other port 1092 connects to the low pressure side of the air piston 1034 through passage 1096 and through a No. 56 bleed 1098 to the body static ports 130 through passage 1100 and tube 1038. At low flight Mach numbers, an upper land 1102 of the shorting valve 1030 is covering the ports 1092 leading to the low pressure side of the air piston 1034. The inlet ports 1090, from the high pressure side of the air piston 1034, are never covered by the valve spool 1076. By venting the high side pressure of the air piston 1034 to the inside ends of both lands 1102 and 1104, pressure balance on the spool 1076 is maintained.

As the Mach sensing bellows 1024 sees a Mach number slightly below a value of 1.6, it free end 1028 has moved enough to allow the shorting valve 1030 to be on the verge of uncovering the sleeve port 1092 to the low pressure side of the air piston 1034. As the Mach number, as seen by the Mach sensing bellows 1024, increases slightly, the shorting valve 1030 begins to open and reduce the pressure on the high pressure side of the air piston 1034, venting it to the low pressure side of the air piston and to atmosphere. When the Mach sensing bellows 1024 sees a Mach number equal to 1.6, the differential pressure (high pressure minus low pressure) of the air piston 1034 has decreased by about one third. When the Mach sensing bellows 1024 sees a Mach number equal to 1.6, the velocity of the missile 100 should be 2,000 ft/sec. As the air piston shorting valve 1030 begins to open, the sensitivity equalizer 1052, which is located on the top of the Mach bellows housing 1066, comes into action as will be explained below. When the shorting valve 1030 opens further, the differential pressure of the air piston 1034 drops to zero.

The air piston 1034 is of the free floating type. A corrected ram pressure from the spillover bleed circuit is fed by tube 872 to the high pressure side of the air piston 1034 and atmospheric pressure by passage 1096 to the low pressure side of the air piston 1034. The force imposed on the air piston 1034 by this differential pressure is applied to the fuel piston 1054 by means of a force pin 1106 which passes through the body casting of the air piston 1034.

The lower side of the fuel piston disc 1054 is connected to the fuel poppet valve 1058 by a fuel piston rod 1108. The upper side of the fuel piston 1054 has forces applied to it by means of the force pin 1106 from the air piston 1034. Partially metered fuel pressure is fed from the fuel flow straightener 1056 to the high pressure side of the fuel piston 1054 through a tube 1110 having apertures 1112 therein. Fully metered fuel pressure is fed to the low pressure side of the fuel piston 1054 through a passage 1114 connected to a tube 1116 having a pressure pick-up point 1118 located in the outlet 1120 of the main fuel regulator 380.

The force imposed on the fuel piston 1054 by this differential pressure is opposed to the force applied by the differential pressure of the air piston 1034. These forces are equal when the system is in a stable metering state. A change in the force balance due to a change in the air heads, or the fuel pressures, causes the fuel metering linkage (air piston 1034 and force pin 1106, fuel piston 1054 and piston rod 1108, and fuel poppet valve 1058) to move in such a manner as to restore the force balance. Thus, a drop in air piston differential pressure would cause the entire linkage to move upward toward the closed position of the fuel poppet valve 1058 until a new equilibrium point is reached.

The fuel flow poppet valve 1058 is of the stainless steel piston type. It is connected to and actuated by the fuel piston 1054. It fits into a poppet valve stainless steel sleeve (not shown) which, in turn, fits the fuel inlet side of the regulator body casting. It is sealed at its ends by O rings. This is the first point of the fuel flow metering and it will control the differential pressure across the variable area orifice valve 1064 as determined by the differential pressure of the air piston 1034 (missile velocity). When the poppet valve 1058 moves toward the Mach sensing bellows 1024, the fuel flow is reduced. Four ports (not shown) are cut around the sleeve to permit fuel flow through the sleeve.

The bottom end 1122 of the poppet valve 1058 is blanked off except for two holes 1124 and 1126. The hole 1124 in the center of the valve 1058 is used for attaching the valve to the valve rod 1108; while the second hole 1126 is utilized to equalize the pressures above and below the blanked end 1122. Slots (not shown) can be provided in the bottom of the valve 1058 to reduce weight and G forces during the flight of the missile 100. These slots are cut in the valve 1058 below the level of the lower edges of the fuel ports in the valve sleeve (not shown) when the poppet valve 1058 is closed.

In practice, a sleeve cover holds the sleeve in place in the regulator body by means of lock wired screws, and is sealed against leakage by an O ring. The valve rod 1108 is attached to the valve 1058 by means of a nut 1128 which engages the threaded end of the rod 1108. A coil spring 1130, together with a flange ring 1132 mounted integral to the valve rod 1108, is utilized to bias the valve 1058.

The fuel flow straightener 1056, FIG. 10, is located between the fuel poppet valve 1058 and the fuel piston 1054. The fuel, after passing the fuel poppet valve 1058, first passes through the straightening baffle 1056 and enters the center of a second baffle (not shown). The second baffle is comprised of two perforated spaced cylinders.

The pressure pick-up to the high pressure side of the fuel piston 1054, as previously mentioned, is through the series of five holes 1112 (only two of which are shown) drilled in the tube 1110. The tube 1110 is mounted on the side of the fuel gathering annulus 1134 away from the fuel inlet to the variable area altitude valve 1064. The series of two baffles mentioned above reduce turbulance and local velocity heads created by turbulance, and gives an essentially true and steady pressure to the fuel piston 1054.

The sensitivity equalizer 1052, previously mentioned, is comprised of the small piston 1032 fed on one side through passage 1136 by air pressure from the high pressure side of the air piston 1034 and on the other side by pressure from a separate air bleed circuit 1036. The forces from the sensitivity equalizer piston 1032 act on the free end 1028 of the Mach sensing bellows 1024 through the force pin 1138. The purpose of the sensitivity equalizer 1052 is to equalize the sensitivity of velocity control mechanism 1044 with changes in altitude by preventing excessive travel of the shorting valve 1030 with a given velocity overshoot at low altitude (high air head), thereby permitting use of a sensitive high altitude (low air head) bellows. This results in increased sensitivity at high altitudes while holding a given maximum sea level sensitivity.

In addition to the small piston 1032, the sensitivity equalizer 1052 is comprised of a spring 1140, a force pin 1138, and a piston sleeve 1142. The bottom of the piston 1032 is fed by air pressure from the high side pressure of the air piston 1034, and the top by the previously mentioned special bleed circuit 1036 and the spring 1140. The force pin 1138 connects the bottom of the piston 1032 and the free end 1028 of the Mach sensing bellows 1024.

Before the shorting valve 1030 opens, the air pressure acting on the bottom side of the sensitivity equalizer piston 1032 is higher than that operating on the top of the piston. However, the force pin 1138 is held against the free end 1028 of the Mach sensing bellows 1024 by the force of the spring 1140 acting through the piston 1032. The entire shorting valve 1030, Mach sensing bellows 1024, and sensitivity equalizer assembly 1052 is supported by the two springs 1080 and 1140 at the ends thereof. All spring forces, including the spring force of the Mach sensing bellows 1024 are cancelled out so that the spring rates of both springs 1080 and 1140 and the Mach sensing bellows 1024 are added.

At low altitude, as the velocity of the missile 100 approaches 2,000 ft/sec from below, the Mach control mechanism 1044 and sensitivity equalizer 1052 are in a state of rest with all forces balanced and the shorting valve 1030 is closed. As the missile 100 enters the speed control range, the shorting valve 1030 opens slightly, dropping the pressure rapidly on the high pressure side of the air piston 1034. As this pressure is dropped, the pressure on the bottom of the sensitivity equalizer piston 1032 is also dropped. The pressure on the top of the piston 1032, combined with the spring pressure on the piston, tends to oppose further motion of the Mach sensing bellows 1024 and shorting valve 1030 and it reduces the travel of the shorting valve 1030 for a given increase in the ram-pressure atmospheric pressure ratio.

As the ram pressure atmospheric pressure ratio (Mach number) increases still further, the Mach sensing bellows 1024 opens the shorting valve 1030 still more, thus reducing the pressure on the high side of the air piston 1034 and the bottom of the sensitivity equalizer piston 1032 even more. The sensitivity equalizer spring 1140 and the pressure on the top of the piston 1032 oppose further opening of the shorting valve, with the rate of opening for the shorting valve, for a given increase in ram pressure atmospheric pressure ratio, being reduced even further. At velocities high above the speed control range, with the pressure on the bottom of the sensitivity equalizer piston 1032 at atmospheric (shorting valve 1030 open), the sensitivity equalizer 1052 applies a nearly constant force to the Mach sensing bellows 1024 and its effect is essentially the same as a higher spring rate in the Mach sensing bellows.

The result is that the sensitivity equalizer 1052 is out of the system, and the rate of travel of the Mach sensing bellows 1024 and shorting valve 1030, for a given increase of ram pressure atmospheric pressure ratio, is nearly the same as without the sensitivity equalizer 1052 present. At velocities far below the design speed range (shorting valve 1030 full closed), the sensitivity equalizer 1052 does not affect the travel of the Mach sensing bellows 1024 for a given increase in ram pressure atmospheric pressure ratio. At high altitudes, when air pressures are low, the effect of the sensitivity equalizer 1052 is much reduced, and the travel of the Mach sensing bellows 1024 is dependent primarily upon the pressure ratio applied to the Mach sensing bellows.

The velocity sensing mechanism 1072 is mounted on a housing 1144 cast integral with the air piston assembly cover 1088. The bellows 1014 is secured at one end 1146, of the housing 1144, and it is sealed by a brass plate soldered to the free end 1148 of the bellows 1014. The free end 1148 of the bellows 1014 is connected to a variable area outlet bleed 1012 by a piano wire 1150, and it positions the variable area outlet bleed 1012 in accordance with differential pressure signals generated by the previously mentioned temperature pick-up 1016 mounted in the ramjet diffuser 270. Diffuser stagnation pressure is applied to the outside of the bellows 1014, and a pressure generated by the bi-metallic controlled poppet valve 1020 and the downstream bleed 1022 is applied to the inner side of the bellows 1014.

The variable outlet bleed 1012 comprises a spool type valve 1152 and a sleeve 1154 in which it operates. The spool valve 1152 is powered by the velocity sensing bellows 1014 by means of the piano wire 1150 linked to the bellows 1014. The pressure from the bi-metallic controlled poppet valve 1020 is led to each end of the spool 1152 to eliminate pressure unbalance on the valve ends. A corrected ram pressure is fed by tubes 868, 870, 872, and 1156 to the center of the spool valve 1152. This pressure acts on both inner faces of the spool valve to eliminate any pressure unbalance. The sleeve 1154 fits inside the velocity assembly casting 1144, which is integral with the air piston housing cover. The sleeve 1154 is sealed against leakage within the housing by O rings 1158.

The sleeve 1154 contains two sets of ports 1160 and 1162. The inlet port 1160 consists of four radially drilled holes fed from an annulus 1164 in the velocity assembly housing 1144. Ram pressure is fed to the annulus 1164 surrounding the sleeve inlet ports 1160 by means of the No. 65 bleed 1010, and from the annulus 1164 to the inner bellows 1035 of the Mach sensing bellows 1024. The outlet ports 1162 of the sleeve 1154 consists of two slots. These slots are cut in the sleeve 1154 in the same direction as the travel of the valve spool 1152. These slots 1162 feed an annulus 1166 in the housing 1144 which is vented by passages 1168 and 1100 to static pressure. The design of the spool valve 1152, sleeve 1154, and ports 1160 and 1162 is such that the inlet ports 1160 are never covered by one end of the spool valve 1152, while the other end of the spool valve 1152 varies the area of the outlet ports 1162 with the travel of the valve.

Ram air pressure is fed through the inlet ports 1160 of the sleeve 1154 to the spool valve 1152 between the lands of the valve and hence to the inner bellows 1035 of the Mach sensing bellows 1024. The ram air is bled off between the lands of the velocity control spool valve 1152 and from the inner bellows 1035 of the Mach sensing bellows 1024 at a rate dependent upon the position of the spool valve 1152, which determines the amount that the outlet port 1162 is open. The position of the valve 1152 is, in turn, dependent upon the pressure differential developed by the temperature pick-up 1016, which is applied to the velocity sensing bellows 1014.

Thus, on increasing the velocity, at a constant altitude, the temperature is increased. Higher temperatures increase the pressure ratio developed by the temperature sensing pick-up 1016 and applied to the velocity sensing bellows 1014. Increased pressure applied to the velocity sensing bellows 1014 moves the bellows and spool valve 1152 in a direction to close the variable area bleed 1012. Closing the bleed 1012 reduces the amount of air bled off from between the lands of the spool 1152 and the inner bellows 1035 of the Mach sensing bellows 1024. Since the air flow through the No. 65 inlet bleed 1010 is essentially constant, reduction of the amount of air bled off by the variable area bleed 1012 increases the pressure applied to the inner bellows 1035 of the Mach sensing bellows 1024.

At constant velocity and constant altitude, increasing the air temperature increases the pressure ratio produced by the temperature sensing pick-up 1016 and applied to the velocity sensing bellows 1014. Increasing the pressure applied to the velocity sensing bellows 1014 causes the bellows and the spool valve 1152 to move so as to reduce the area of the variable area bleed 1012, thus reducing the amount of air bled off from between the lands of the spool valve 1152 and the inner bellows 1035 of the Mach sensing bellows 1024. At constant velocity, an increase in ambient temperature results in a decrease in Mach number. The action of the velocity control system is to prevent this decrease in Mach number from being seen by the Mach sensing bellows 1024 while the missile 100 is flying at a constant velocity.

The temperature pick-up 1016 shown in FIGS. 9A and 73, while not physically a part of the main fuel regulator 380, is discussed with the main fuel regulator 380 as its operation is closely tied into the operation of the main regulator. The temperature pick-up housing 1168 is flanged to mount on the wall of the diffuser 270, drilled for a series of bleeds 1022, 1170, and 1172, and the seat fitting 1174 of the poppet valve 1020. A portion of the housing 1168, projecting into the air stream in the diffuser 270 is ducted, by duct 1176, to reduce the velocity head of the air flowing through it to essentially zero, resulting in maximum temperature and pressure recovery at the temperature pick-up 1016. The location of the predetermined temperature pick-up 1016 in the diffuser 270 functions minimize error in stagnation temperature measurement. The temperature pick-up 1016 will operate satisfactorily at stagnation temperatures ranging from 160.degree. to 530.degree. F, and diffuser stagnation pressures ranging from 6 psia to 95 psia. The differential pressure generated will be a function of stagnation temperature.

The temperature sensing element 1018 in the temperature pick-up 1016 is a high response bi-metallic strip. The strip 1018 is mounted as a cantilever beam. A fixed end 1178 of the beam 1018 is clamped into a mounting machined in the duct of the temperature pick-up 1016. A free end 1180 of the beam 1018 has a slotted hole cut near its edge, into which fits the upper end of a pick-up poppet valve stem 1182.

The poppet valve seat 1174 is screwed into a hole 1184 in the pick-up housing 1168 directly below the end of the bi-metallic strip 1018. It is secured in place by a lock nut 1186. Adjustment of the pressure ratio acting across the velocity bellows 1014 is made by adjusting the position of the pick-up poppet valve seat 1174. The poppet valve 1020 is a light weight, cone shaped valve connected to the bi-metallic 1018 by the piano wire 1182. A fine mesh screen 1188 is located immediately below the poppet valve 1020 to prevent clogging of the bleeds in the velocity control bleed circuit. The outlet bleed 1172, in series with the poppet valve 1020, is vented to static pressure. The bleed 1022 is placed in the line from the poppet valve outlet to act as a surge preventer and permits smoother operation of the unit.

The pressure pick-up feeding the high pressure side of the velocity bellows 1014 by means of line 1190 also contains surge preventing bleed 1170. The line 1190 is a pressure transmission line only; and the bleed 1170 in this line 1190 is merely a restriction to any rapid pressure changes.

The bi-metallic element 1018 is so mounted that increasing temperature tends to make the free end 1180 of the bi-metallic beam 1018 curl upwardly. As the free end 1180 of the bi-metallic beam 1018 is secured to the poppet valve 1020 through the piano wire 1182, tension in the piano wire 1182 increases with increasing temperatures. This increased tension in the piano wire 1182 is balanced by the increased pressure drop across the poppet valve 1020 generated by the flow of air through the poppet valve. The design and configuration of the poppet valve 1020, bi-metallic strip or beam 1018, and bleed circuit is such that the pressure differential developed by the temperature pick-up 1016 is relatively independent of altitude (static pressure), diffuser pressure, and is dependent only upon the diffuser stagnation temperature.

The altitude sensing bellows 1060, shown in FIG. 10, is a spring loaded evacuated bellows which has one fixed end 1192 and one free end 1194, which is connected to the variable area orifice valve servo 1062 by a piano wire link 1196. Atmospheric pressure is fed to this bellows 1060, which causes it to position the valve servo 1062. With decreasing pressure (increasing altitude), the bellows 1060 increases in length at the rate of approximately 0.018 inch per inch of mercury pressure decrease.

The variable area orifice valve servo 1062 is a pressure balanced positioning mechanism consisting of a servo pilot valve 1198 and a servo follower 1200. The servo pilot valve 1198 is a pressure balance spool valve, drilled longitudinally to equalize pressures at each end of the spool. The motion of this valve 1198 is a function of altitude, being moved by the piano wire link 1196 connected to the free end 1194 of the altitude sensing bellows 1060. The servo pilot valve 1198 controls the flow of high pressure fuel through the servo follower 1200 to and from the top of the variable area orifice valve servo 1062.

The variable orifice valve servo follower 1200 is a sleeve type follower with three lands on its outside surface. Lands 1202 and 1204 are located at each end of the follower 1200, and the remaining land 1206 is located in the center thereof. These three lands 1202, 1204 and 1206 form two annulus 1208 and 1210 around the servo follower 1200, one of which carries regulator inlet fuel pressure (high pressure) and one of which carries servo operating pressure to the top of the variable area orifice valve 1064.

The servo follower 1200 is connected to the top of the variable area orifice valve 1064 through a flexible linkage 1212. The flexible linkage 1212 is used to prevent any misalignment of components to cause binding in a stable state. The altitude sensing bellows 1060 has caused the pilot valve 1198 to assume a given position. The servo follower 1200 has reacted to assume the equilibrium state, and there is no motion of the system or fluid flow through the assembly.

As the missile 100 climbs to altitude, the altitude sensing bellows 1060 expands and moves the servo pilot valve 1198 toward the variable area orifice valve 1064. As the pilot valve 1198 moves, it uncovers a servo outlet port 1214. As this port 1214 is uncovered, high pressure fuel flows from a high pressure fuel inlet port 1216, through an annulus 1218 between the servo pilot valve 1198 and the servo follower 1200 and thence through the servo outlet port 1214 to the top of the variable area orifice valve 1064.

As fuel flows to the top of the variable area orifice valve 1064, the valve 1064 is moved to the left (closed position) against a spring 1220 and the regulator outlet pressure. The servo follower 1200, being mechanically linked to the variable area orifice valve 1064, moves to cover the servo outlet port 1214. When the port 1214 is closed, fuel flow is cut off. The variable area orifice valve 1064 no longer moves, and a new equilibrium point is established.

When the missile 100 decreases altitude, the altitude sensing bellows 1060 decreases its length, moving the servo pilot valve 1198 away from the variable area orifice valve 1064, and opening the servo outlet port 1214 to the overboard drain 990, releasing fuel pressure from the top of the variable area orifice valve 1064. Spring pressure plus regulator outlet pressure on the bottom of the variable area orifice valve 1064 causes this valve and the servo follower 1200 to move to the right until a new equilibrium point is reached.

The variable area orifice valve 1064 is of a piston type with a series of ports 1222 machined in the valve surface. The valve 1064 is fitted to a sleeve 1224 that fits into the main regulator body casting 1050.

The valve 1064 is spring loaded to the full open position (low altitude position) until acted on by fuel pressure fed by the variable area orifice servo 1062 to oppose the spring force and regulator outlet pressure acting on the bottom of the valve 1064. The port area must be coordinated with the altitude-travel relationship of the evacuated bellows-servo-altitude valve assembly.

The cast aluminum regulator outlet elbow 1120 bolts to the main regulator body casting 1050, and provides support for the variable area orifice spring 1220 and the regulator outlet pressure pickup 1118. The regulator outlet pressure pick-up 1118 consists of a pointed tube mounted parallel to fuel flow from the variable area orifice valve outlet 1120. The tube 1118 has a series of holes 1226 along its length to pick up the regulator outlet static pressure. This pressure is fed to the low pressure side of the fuel piston 1054 through channel 1114 drilled in the mainbody casting 1050, as previously mentioned.

In order to explain the overall operation of the main fuel regulator 380, the following initial conditions will be assumed:

Standard day ______________________________________ Velocity 1,900 feet per second Altitude Sea level Static pressure 14.7 pounds per square inch Static temperature 59.degree. F Mach number 1.702 Ram pressure 62.22 pounds per square inch Ram to static pressure ratio 4.2325 Stagnation temperature 359.3.degree. F Temperature pick-up .DELTA.P 7.2 inches Hg. ______________________________________

During the above conditions, the temperature pick-up 1016 impresses a differential pressure of 7.2 inches of mercury across the velocity sensing bellows 1014, causing it to move the variable area bleed 1012 to such a position as to bleed sufficient pressure from the inner bellows 1035 of the Mach sensing bellows 1024 to lower the pressure ratio acting across the Mach sensing bellows 1024 to below a Mach number equal to 1.6 with a pressure ratio of ram to static of 3.8. The Mach sensing bellows 1024 has, however, moved the air piston shorting valve 1030 to a position such that it is near the position where it will open. The main regulator air piston differential pressure .DELTA. P is such under these conditions to make the main fuel flow fully rich, giving the missile 100 excess thrust and causing it to accelerate.

If the missile 100 now accelerates to 2,000 ft/sec with no altitude or ambient temperature changes, the following conditions exist:

Velocity 2,000 feet per second Altitude Sea level Static pressure 14.7 pounds per square inch Static temperature 59.degree. F Mach number 1.792 Ram pressure 68.11 pounds per square inch Ram to static pressure ratio 4.633 Stagnation temperature 391.8.degree. F Temperature pick-up .DELTA.P 7.8 inches Hg.

As the missile velocity increases, the pressure differential generated by the temperature pick-up 1016 and applied across the velocity sensing bellows 1014 increases, causing the velocity sensing bellows 1014 to close the variable area bleed 1012. This decreases the amount of pressure bled off the inner bellows 1035 of the Mach sensing bellows 1024. At the same time, the ram pressure is increasing, thus increasing the pressure applied to the inner bellows 1035 of the Mach sensing bellows 1024 through the velocity control inlet bleed 1010.

As the pressure applied to inner bellows 1035 of the Mach sensing bellows 1024 increases, the pressure ratio across the Mach sensing bellows 1024 increases since the static pressure applied to the outside of this Mach sensing bellows remains constant. At this pressure ratio increases, the Mach sensing bellows 1024 expands, moving the air piston shorting valve 1030 toward an open position. When the pressure ratio applied across the Mach sensing bellows 1024 reaches 3.8 (pressure ratio at a Mach number equal to 1.6), the shorting valve 1030 is on the verge of opening and any increase in pressure ratio opens the air piston shorting valve 1030 and shorts the high pressure side of the air piston 1034 to the low pressure side.

As the air piston differential pressure is decreased, the fuel metering linkage moves to reduce the flow of fuel through the main fuel regulator 380. On reducing the fuel flow to the combustor 108, there results a decrease in thrust so that the missile 100 no longer accelerates. Under normal flight conditions, when the main fuel regulator 380 goes lean (fuel flow reduces), this regulator will act in such a way that main fuel flow will be at some intermediate point between full rich and full lean, and sufficient thrust from the main combustor 376 will be provided to maintain the flight velocity of the missile 100 at 2,000 ft/sec.

With the missile 100 flying at sea level at a velocity of 2,000 ft/sec during a standard day, let it now be assumed that flight conditions are instantaneously changed to those of a proposed military cold day. For a proposed military cold day, flight conditions will be assumed as follows:

Altitude Sea level Ambient pressure 14.7 pounds per square inch Ambient temperature -22.degree. F Flight velocity 2,000 feet per second Mach number 1.951 Stagnation temperature 310.degree. F Temperature pick-up .DELTA.P 6.2 inches Hg. Ram to static pressure ratio 5.3929 Ram pressure 79.3 pounds per square inch

It will be noted that for a cold day, with a flight velocity of 2,000 ft/sec, the flight Mach number as seen by the missile 100 is increased. This is due to the fact the velocity of sound in the atmosphere is dependent upon ambient temperature. In fact, the velocity of sound in the atmosphere is proportional to the square root of the absolute ambient temperature. As the Mach number is defined as the velocity of the air stream moving past an object (or the velocity of the object moving through the air), divided by the sonic velocity in undisturbed air, it can be seen that for a constant flight velocity, if the ambient air temperature is changed, that the Mach number will change.

At the new conditions, a larger mass of air flows through the combustor 108. Denser air flows past the outside of the missile 100, which results in an increase in drag. The main fuel regulator 380 acts in such a manner as to increase the fuel flow necessary to provide an increase in thrust to overcome the increase in drag and maintain velocity of the missile 100 constant. The decrease in ambient temperature results in a decrease in stagnation temperature at flight conditions. This decrease in stagnation temperature causes the bi-metallic element 1018 to exert less force on the pick-up poppet valve 1020, which results in less pressure drop across the poppet valve 1020 and a lower pressure differential applied to the velocity sensing bellows 1014.

The lower pressure differential applied across the velocity sensing bellows 1014 causes the variable area bleed 1012 to open further, thus bleeding off more of the ram pressure applied to the Mach sensing bellows 1024. This results in the Mach sensing bellows 1024 seeing no change in applied pressure differential across it. The air piston shorting valve 1030 does not open further. An increase in ram pressure applied to the air piston inlet bleed 1228 results in a higher air piston differential pressure being applied to the air piston 1034, which moves the fuel metering linkage in such a manner to cause an increase in pressure drop across the variable area orifice valve 1064, resulting in an increase in fuel flow.

Let it now be assumed that the missile 100 is instantaneously raised to 40,000 feet altitude, on a cold day, with a flight velocity of 2,000 ft/sec. The flight conditions will now be assumed as follows:

Altitude 40,000 feet Ambient pressure 2.722 pounds per square inch Ambient temperature -103.degree. F Flight velocity 2,000 feet per second Mach number 2.160 Stagnation temperature 229.8.degree. F Temperature pick-up .DELTA.P 5.2 inches Hg. Ram to static pressure ratio 6.4931 Ram pressure 17.67 pounds per square inch

The stagnation temperature change with altitude results in a lower temperature pick-up .DELTA. P being applied across the velocity sensing bellows 1014, opening the variable bleed 1012 further to bleed off more of the ram pressure applied to the Mach sensing bellows 1024, and lowering the pressure ratio seen by the Mach sensing bellows 1024 to its previous value. The lower ram pressure applied to the air piston inlet bleed 1228 results in a lower air piston .DELTA. P, which causes the fuel metering linkage to move in such a manner as to reduce the pressure drop across the variable area orifice valve 1064.

With the increase in altitude, the static pressure applied to the altitude sensing bellows 1060 has decreased. This decrease in pressure causes the altitude sensing bellows 1060 to increase in length, moving the servo pilot valve 1198 toward the variable area orifice valve 1064. This action vents high pressure fuel to the top of the variable area orifice valve 1064, moving it toward the closed position and moving the servo follower 1200 to close the follower ports 1214, thus reducing fuel flow as a function of altitude. The correct pressure drop across the valve has already been determined by the fuel metering linkage.

The effects of acceleration of the missile 100 are primarily confined to the pilot and main fuel regulators 340 and 380, respectively. Operation of the fuel and nitrogen shut-off valves 428 and 446, respectively, which depend on boost acceleration of the missile 100 for operation has been discussed previously. The effect of acceleration of the missile 100 on the fuel in the inlet line 436 to the pump 532 was considered in establishing the regulated nitrogen pressure.

Minimizing of the acceleration effects on the pilot and main fuel regulators 340 and 380, respectively, is accomplished by aligning the axes of the moving parts parallel to the longitudinal axis of the missile 100 where, after boost of the missile 100 has ended, minimum accelerations are expected. Furthermore, the design of the main fuel regulator 380 is such that longitudinal acceleration effects on the altitude valve control mechanisms cause an error in the opposite sense from the corresponding error induced in the pressure regulating linkages. This cancellation effect also occurs between the double bellows 1033 and 1035 of the Mach sensing bellows 1024 and the temperature controlled bellows linkages in the velocity control mechanism 1044.

Squib activated time delay units 866 and 1230, as shown in FIGS. 9A and 70, may be used in the missile 100 to lower the fuel flows of the pilot fuel regulator 340 at separation of the missile 100 from the booster 500 to insure smooth ignition of the air-fuel mixture, and to lock the high pressure side of the main fuel regulator 380 to the low side before separation of the missile 100 from the booster 500 to reduce the main fuel flow to a minimum during boost of the missile 100, respectively.

The squib activated time delay unit, such as 866 or 1230, is a two position valve actuated by an electrically fired squib 1006. These units are commonly used to control air flows and to change air pressures applied to various components in a system.

The squib activated time delay unit 866 or 1230 is comprised of a housing 988, a piston valve 1002 so drilled as to permit free passage of air from near one end of the piston 1002 to the other end, a heavy spring 1232, a housing cover 1234 which holds the spring 1232 in place, a piston actuating rod 1236, and an actuating rod lever 1238. A squib activated time delay unit 866 or 1230 in the "unfired position" is shown in FIG. 70.

In the unfired position, the piston valve 1002 rides in a hole 1240 drilled through a housing partition 1242. The piston valve 1002 is held by the actuating rod 1236 and actuating rod lever 1238 against the pressure of the spring 1232 so that air is free to flow from one end of the piston 1002 to the other by passage 1244 drilled in the piston 1002. The end of the piston 1002, upon which the spring 1232 acts, is of a larger diameter than the piston proper.

An O ring 1246 is placed around the piston 1002 beneath the enlarged end 1248 of the piston to seal against air leakage when the piston valve 1002 is closed. When the unit 866 or 1230 is in the unfired position, the actuating rod lever 1238 is lock wired to the body by means of a soft copper wire 1004.

This wire 1004 extends through a chamber 1250 into which the squibs 1006 project and in which they fire. When the actuating rod lever 1238 is lock wired shut, it forces the actuating rod 1236 into the housing 988, and opens the piston valve 1002 against the pressure of the spring 1232. When a squib 1006, mounted in the squib activated time delay unit 866 or 1230, is fired, the soft copper wire 1004 is severed, releasing the actuating rod lever 1238. The pressure of the spring 1232 against the piston valve 1002 forces the piston valve to the closed position against friction of the actuating rod 1236. As the piston valve 1002 closes, the O ring 1246 seals against leakage around the hole 1240 in the partition 1242.

The squib 1006, used to activate these units 866 or 1230, is electrically fired, and contains a small powder charge which shatters a copper cap (not shown) surrounding the squib 1006 upon firing. These squibs 1006 are potentially dangerous, and their two leads are grounded to each other when not installed in the missile 100. A 4.5 volt battery supplies enough current to fire these units 866 or 1230.

In the event it is not desired to use the squib activated time delay unit 1230, it may be taken out of the main regulator air regulation circuit. It may be used to control the flow of regulated nitrogen pressure to the pilot regulator 340.

Fuel flowmeters 460 and 462 can be incorporated in the fuel system of missile 100 to measure fuel flow through the pilot and main fuel regulators 340 and 380, respectively, in order to determine proper operation of the missile fuel system. Information measured by these flowmeters 460 and 462 is fed into the missile telemetering system for transmission to the ground during aerial flight of the missile 100.

The main fuel flowmeter 462 is mounted in the fuel outlet 1120 from the main regulator 380 and it measures fuel flow through the main regulator, while the pilot fuel flowmeter 460 is mounted in an outlet fuel line 1252 from the pilot regulator 340 and it measures the fuel flow through the pilot regulator.

The main and pilot flowmeters 462 and 460 are similar, with individual differences being principally in size of the bore which affect fuel flow versus frequency. As can best be seen in FIG. 72, each flowmeter 460 and 462 consists of a body casting 1254 containing a flowmeter bore 1256, for fuel flow, a flow straightener (not shown), jeweled mountings 1258 and 1260, a rotor-shaft assembly 1262, a photoelectric cell 1264, and a light source 1266. A comparison of the bore sizes and frequency characteristics of these flowmeters is given below:

Main Fuel Pilot Fuel Flowmeter Flowmeter ______________________________________ Bore size, diameter 1.235 inches 0.896 inches Fuel flow range 0-6.5 pounds 0-4.25 pounds per second per second Frequency range 0-160 cps 0-245 cps ______________________________________

The bore 1256 and bearing supports 1258 and 1260 are streamlined to provide smooth flow through the flowmeter 460 or 462. The straightener (not shown) is fitted to the inlet of the flowmeter 460 or 462 to insure smooth fuel flow therethrough.

The rotor-shaft assembly 1262 consists of two vanes 1268 mounted on a shaft 1270. The entire assembly 1262 is blackened except for a small polished insert 1272 at one end of the shaft 1270. This polished insert 1272 is wrapped around the rotor-shaft 1270, and is so shaped that as the shaft 1270 rotates, the area passing under a given point varies sinusoidally at the frequency of the shaft rotation. The light source 1266 is so placed in the body casting 1254 as to shine on the polished insert 1272. The light is then reflected to the photocell 1264 mounted in the same plane as the light source. The light source is a 4.5 volt 0.36 amp DC silica bulb. The bulb 1266 is mounted behind a sealed lens (not shown) which is so mounted as to cause the light from the bulb 1266 to shine on the polished insert 1272 in the shaft 1270.

The photocell 1264 used in the flowmeters consists of a metal plate on which is deposited a thin layer of light-sensitive, conductive selenium. On top of this light sensitive layer of selenium is deposited a transparent layer of a special alloy. The iron base plate is sprayed with a non-corrosive metal and the entire unit is sealed with lacquer. This cell 1264 results in a direct transformation of light energy into electrical energy. The current produced varies directly as the intensity of the light reaching the photocell 1264. The photocell 1264 is mounted in the flowmeter 460 or 462 behind a transparent lens (not shown) which is sealed against fuel leakage by an O ring. In practice, the photocell 1264 is held in place by a brass plunger fitting in a plastic screw plug. The brass plunger provides electrical contact with the photocell 1264 and mechanically holds it in contact with the lens.

In operation, fuel flow through the bore 1256 of the flowmeter 460 or 462 causes the rotor 1262 to rotate at a velocity dependent on the fuel flow rate. The light source 1266 delivers a light of constant intensity to the rotor insert 1272. The constantly varying area of the insert 1272 presented to the light 1266 by the rotation of the shaft 1270 causes light of varying intensity to be reflected to the photocell 1264, whose output varies in intensity. The photocell 1264 output is, therefore, a sinusoidally varying voltage, the frequency of which depends upon the rotational velocity of the shaft 1270.

The output voltage of these flowmeters 460 and 462 is on the order of 10 to 30 millivolts. The input to each light bulb 1266 is 4.5 volts (DC) and 350 milliamperes. The expected frequency operating range for both fuel flowmeters 460 and 462, respectively, is 10 to 250 cps.

The output frequency of the flowmeters 460 and 462 is directly proportional, with a high degree of accuracy, to the volume of fuel flowing. Frequency type measurements insure that noise and attenuation accompanying the telemetering signals during flight of the missile 100 do not destroy the accuracy of the readings. Both flowmeters 460 and 462 are constructed to be replaced by simple flow connectors in the event it is determined that fuel flow measurements are no longer deemed necessary.

The probe static relief valve 1048 is installed in the probe static line 1046 at the nose of the missile 100 to prevent damage to the Mach sensing bellows 1024 by accidental high probe static pressures.

This valve 1048, FIG. 71, is a ball type, spring loaded check valve, designed to open at between 4.5 psi and 7.5 psi above compartment pressure. The valve 1048 is installed to relieve any high pressure transients imposed upon the probe static line 1046 during the flight of the missile 100 and to prevent these high pressure transients from damaging the Mach sensing bellows 1024. High pressure transients will be imposed on the probe static line 1046 whenever an oblique or plane shock wave passes over the probe static inlet 126.

A T-fitting 1274 is installed in the line 502 between the nitrogen relief valve 452 and the inlet fitting 504 for the bladders 438, as seen in FIG. 9B. One side of this fitting 1274 leads to the turbine scoop 772. The check valve 450 is positioned in the line 502 to allow free flow of air from the turbine scoop 772 to the bladders 438 when the scoop pressure is higher than the pressure in the nitrogen system. This check valve 450 is spring loaded to prevent flow of nitrogen from the nitrogen system to the turbine scoop 772. This air connection, by lines 528 and 502, permits air at the pressure in the diffuser 270 to pressurize the bladders 438 whenever the nitrogen system pressure falls below pressure in the diffuser 270. For example, the pressure in the diffuser 270 would be higher than the nitrogen pressure during a low altitude intercept of the missile 100 after a long range "up and over" flight.

It is important that this check valve 450 be placed in the missile 100 with the valve body perpendicular to the principal axis of the missile 100, with the hinge pin forward. When this valve 450 is in this position, acceleration forces during launching of the missile 100 will help hold the check valve 450 closed, and prevent loss of nitrogen.

Fuel injectors are provided as part of the fuel system to introduce the fuel into the air stream in a finely divided state, well mixed with the air to insure rapid and complete burning thereof.

As illustrated in FIGS. 9A, 31, and 32, fuel leaving the main and pilot regulators 380 and 340 first passes through the fuel flowmeters 462 and 460, respectively, in the fuel passages 1276 and 1252 to the turbine air intake strut 424. The air intake strut 424 is secured diagonally across the diffuser 270, in the forward portion of the center body 106 of the missile 100. Channels 1278 and 1280 in the forward side of the strut conduct air from the diffuser 270 to the fuel turbine pump and hydraulic turbine 440 and 1282, respectively. The aft side of the strut 424 to the fuel turbine pump 440 is divided into two channels 422 and 420 which carry main and pilot fuel to the fuel delivery tubes 1276 and 1252, respectively.

The fuel delivery tubes 164 and 166 are concentric tubes mounted on the center line or longitudinal axis of the missile 100. The forward ends of these tubes 164 and 166 are connected with the passages 420 and 422 of the air intake strut 424. The pilot and main fuel delivery tubes 164 and 166 are made of aluminum. The main fuel delivery tube 166 is mounted inside the pilot fuel delivery tube 164 and is arranged concentric with it. The forward ends of both tubes 164 and 166 fit into the aft side of the air intake strut 424 and are sealed with 0 rings (not shown).

The pilot fuel tube 164 terminates in the forward end of a pilot manifold hub 1284. The terminal ends of both tubes 164 and 166 are sealed with O rings 1286 and 1288, respectively. The tubes 164 and 166 are held in place longitudinally by the pilot manifold hub 1284 which is supported by shroud assembly 332 and the air intake strut 424 which is supported in the wall of the diffuser 270. The aft end of the pilot manifold hub 1284 is tapered to a reduced diameter and extends to the forward end of the main fuel distribution valve 384, where it terminates and is sealed by an O ring 1290, shown in FIG. 34.

As seen in FIG. 32, four struts 342 of aluminum tubing are utilized to secure the pilot injector manifold ring 344 to the pilot manifold hub 1284. These struts 342 are spaced 90.degree. apart around the hub 1284 and are welded thereto. The manifold injector ring 344, as previously mentioned, is made of aluminum tubing formed in a circle, drilled for the entry of the struts 342, and it is welded thereto. Sixteen fittings 1292, spaced 221/2 .degree. apart, are welded to the forward side of the injector ring 344, for mounting of the fuel injectors 330. The forward end of these fittings 1292 are threaded so that the pilot injectors 330 can be screwed and lock wired thereon.

In FIG. 74, each of the pilot fuel injectors 330 proper is of a spring loaded pintle type. These injectors 330 are screwed directly to the injector fittings 1292 which, as previously indicated, are welded to the forward side of the manifold ring 344. The body of each injector 330 is formed of a cylindrical piece of stainless steel.

One end of the body of each injector 330 is tapped with threads 1298 for securing it to the injector fittings 1292 on the manifold ring 344. A sleeve 1294, with a shoulder 1296, is pressed into the body at the end opposite the threads 1298. The outer face of the sleeve 1294 is machined with a seat 1300 for a pintle valve 1302. The sleeve 1294 has holes 1304 drilled from its outside to inside surfaces to permit passage of fuel through the sleeve 1294. This sleeve 1294 serves as a guide for the pintle valve 1302; a support point for a spring 1306 which loads the pintle valve 1302; and as the seat for the pintle valve 1302.

The pintle valve 1302 is made of stainless steel and includes a rod threaded at one end. The opposite end of the pintle valve 1302 is machined with a valve disc 1308, which tapers to a point at the extreme end. The pintle valve 1302 has two lands 1310 machined on it which guide the pintle valve 1302 in the sleeve 1294. The pintle valve 1302 is inserted into the valve sleeve 1294 and held in place by the load of spring 1306 acting against a shouldered adjustment nut 1312 screwed to the end of the shaft 1314. The shoulder nut 1312 is locked in place by a small lock nut 1316.

The pressure drop across the injector 330 is adjusted by adjusting the spring force on the pintle valve 1302 by means of the shoulder nut 1312. It is to be noted that the units are designed so that fuel flow rates, at given pressure drops, for each injector 330 must agree within .+-. 5 percent. The pressure drop across the injector 330 varies from 15 psi to 160 psi, depending on the rate of fuel flow. When the injector 330 is passing fuel, fuel pressure, acting on the inner face of the valve disc 1308, causes the pintle valve 1302 to open. As the fuel leaves the injector 330, it is broken up into a very fine, conical, fog like spray. Thus, the pilot fuel is injected against the air stream by the sixteen injectors 330 mounted on the injector ring 344. Combustion of the pilot fuel is initiated by means of the previously mentioned modified aircraft type spark plug or igniter 1008 which is supplied with high voltage by a conventional vibrator-induction coil combination.

As previously mentioned, the aft end of the pilot manifold hub 1284 tapers to a reduced diameter and extends aft for a fixed distance and then terminates in the main fuel distribution valve 384. It is sealed by the 0 ring 1290.

In FIGS. 31, 32, and 35, the main fuel distribution valve 384 is a sliding piston variable area orifice valve made of stainless steel. It has a valve sleeve 1318, with four flats 1320 spaced 90.degree. apart and milled along its outside surface. The aft end of hub 1284 fits into the forward end of sleeve 1318. Four holes 1322 are drilled through the sleeve 1318 on the centerline of each flat 1320.

The moving part of the valve 384 consists of a piston 1324 which is closed at one end. Four rows of holes, each row containing four holes 1326, are drilled through the piston 1324 to match with the four rows of holes 1322 drilled through valve sleeve 1318. A slot (not shown) machined to cooperate with a pin (not shown) inserted in the sleeve 1318 prevents the piston 1324 from rotating and causing misalignment of the ports 1326 and 1322 in the piston 1324 and sleeve 1318, respectively. The closed end of the piston 1324 bears against a valve spring 1328, which is housed in the aft section 1330 of the valve 384. The aft section 1330 contains an adjustable stop 1332 which positions the piston 1324 in its open position, and a back pressure vent 1334 connected to one of the holes 1322 in the valve sleeve 1318 to vent the pressure of the injectors 386 to the back side of the piston 1324.

A fairing 1336 is bolted to the after side of the aft section 1330 to give the valve 384 to clean aerodynamic surface. The piston 1324 is held in the valve sleeve 1318 against spring pressure by a snap ring 1338. When no fuel pressure is applied to valve 384, this valve is held in the closed position against the snap ring 1338 by the valve spring 1328. When fuel pressure is applied to the valve 384 and reaches 22 psi, the piston 1324 has moved enough to cause the ports 1326 in the piston 1324 to uncover the ports 1322 in the sleeve 1318, thereby permitting fuel flow through the ports 1322 to the injectors 386. The maximum pressure drop across the valve 384 is 25 psi.

The fuel from the fuel distribution valve 384 passes to four struts 1340 bolted to the flats 1320 of the valve 384. Each strut 1340 contains four drilled passages 1342 which match with the drilled ports 1322 of the main fuel distribution valve sleeve 1318. After passing through the strut 1340, the fuel enters a main fuel manifold 1344, which contains entrance holes (not shown) matching the passages 1342 in the struts 1340. The struts 1340 and manifold sections 1346 are welded together at their mating faces, as shown at 1348. The main fuel distribution manifold 1344 is manufactured in four 90.degree. sections 1346 from aluminum.

Four equally spaced holes 1350 are drilled in the aft side of each manifold section 1346 for mounting injector tubes 1352. Passages 1354, which connect the holes 1350 with the holes matched to the strut passages 1342, are drilled parallel with the surfaces of the manifold sections 1346 before the manifold 1344 is rolled to circular shape. Both ends of each quarter section 1346 of the manifold 1344 are bolted to a "T" bracket 1356 which secures the manifold 1344 in the forward end 388 of the combustor central can assembly 370.

Further support for the manifold 1344 is provided by the inner ends of the struts 1340 bolting to the flats 1320 on the main fuel distributing valve sleeve 1318. In each of the equally spaced holes 1350 drilled on the aft side of the manifold sections 1346, an aluminum tubing 1352 is welded. The injector end of each tube 1352 is bent outward to form an injection circle. The injector end of each tube 1352 is flared to a reduced diameter for mounting the injectors 386.

The injector body 1358, FIG. 75, is a cylinder. A shoulder 1360 formed by the intersection of the two internal diameters is beveled to form a seat for a piston valve 1362. Two rows of holes 1364 and 1366 are located downstream from this shoulder 1360. Both rows of holes 1364 and 1366 contain six holes which are drilled radially through the injector cylinder 1358. These holes 1364 and 1366 are used for the fuel exit ports. The rows of holes or ports 1364 and 1366 are uncovered, in turn, by the piston valve 1362 operating in the injector cylinder 1358.

Two grooves 1368 and 1370 are machined around the outside of the inlet end of the injector cylinder 1358. The groove 1368, nearest the inlet end of the injector cylinder 1358, contains an O ring (not shown) to seal against fuel leakage between the cylinder 1358 and the injector tube 1352; while the inboard groove 1370 is used to secure the injector 386 to the injector tube 1352.

The injector tube 1352 has two slots (not shown) cut near the end perpendicular to the tube axis which line up with the inboard groove 1370 in the injector cylinder 1358. A lock wire (not shown) wrapped around the end of the tube 1352 engages both the slot and the groove 1370 and secures the injector 386 in place. Both ends of the piston 1362 are beveled to match with the valve seat 1360. The piston 1362 is held in the closed position by an injector spring 1372, which is secured in the injector cylinder 1358 with a snap ring 1374 and spring retaining washer 1376.

The injector spring 1372 is designed to give a pressure drop across the injector 386 of from 10 to 25 psi, depending on fuel flow. The inlet pressure of the injector 386 acts against the piston 1362, and moves it toward the open position, uncovering the row of small ports 1364. A higher fuel flow imposes more pressure on the piston 1364 and moves it to uncover the row of large ports 1366 when high fuel flows are demanded. This low pressure injector 386 is provided in the main fuel system primarily to prevent boiling of the fuel in the injector tubes 1352 and manifold ring 1344 at high altitudes.

As illustrated in FIGS. 15 through 19, the hydraulic system 220 for the missile 100 is an auxiliary power supply used in conjunction with the guidance control system 214. The signal inputs to the hydraulic system 220 consist of differential voltages applied to torque motors 1378, 1379, 1380, and 1381 of the hydraulic system by the guidance control system 214. The output of the hydraulic system 220 is used to operate wing actuators 1382, 1383, 1384, and 1385 which, through the wing linkages, in turn, operate the wings 112, 113, 114, and 115. The hydraulic system 220, together with the guidance control system 214, forms an electro-hydraulic servomechanism to position the pairs of missile wings 112, 114 and 113, 115. The hydraulic system 220 is designed to operate at 1750 psi.

The hydraulic system 220 is a closed loop hydraulic system. It consists primarily of an air turbine 1282 which drives a hydraulic pump 1390, transfer valves 1392, 1393, 1394, and 1395, a sump 1396, and the wing actuators 1382, 1383, 1384, and 1385. Other auxiliary equipment, such as filters, check valves, accumulators, has been added to make the hydraulic system 220 more practicable and workable. A relatively independent sub-system, the wing lock system 1398, has been added to hydraulically unlock the wings 112 through 115 at the separation of the missile 100 from its booster 500, and to signal wind lock retraction to the guidance control system 214 by a pressure operated switch in the wing lock subsystem 1398.

In order to facilitate missile testing, provisions can be made in the hydraulic system to operate the system with an external source of hydraulic power. It is not necessary to disrupt any missile hydraulic plumbing in order to apply external hydraulic power. External hydraulic power can also be applied to the wing lock sub-system 1398. Under these conditions, the wing lock subsystem 1398 is isolated from the remainder of the hydraulic system by check valves.

All hydraulic tubing connecting various components on the high pressure side of the system is constructed of corrosion resistant steel. All low pressure tubing is manufactured of aluminum. Tubing and components are connected by the use of standard fittings.

A hydraulic turbine-pump assembly 1281 is the power source for this system 220. The pump 1390 is driven by the high speed single stage impulse turbine 1282, which is powered by diffuser air from the air intake strut 774. The pump 1390 is driven by the turbine 1282 through a planetary reduction gear assembly 1400 with a reduction ratio of 6.2 to 1. The pump 1390 is a constant displacement axial stroke piston type pump, rated at 20.3 gal/min at a nominal pressure of 1750 psi.

Pump pressure regulation is obtained by means of a movable nozzle plate on the air turbine inlet. This nozzle plate is postioned through a control linkage by the pump outlet pressure, the position of the nozzle plate determining the power input to the pump 1390, speed of the pump and the pump outlet pressure. The throttle control linkage is spring loaded in such a manner as to hold the throttle full open until the pump outlet pressure approaches 1750 psi. When the pump outlet pressure reaches 1750 psi, hydraulic pressure causes the throttle plate to move toward the closed position, reducing the pump speed and the quantity of oil pumped. When an oil demand occurs, a reduction in system pressure takes place, opening the turbine throttle and increasing the pump output.

In order to provide constant pressure in the system, absorb pressure peaks, and provide a reserve of high pressure oil for peak demands, two hydraulic accumulators 122 and 1402 are included in the high pressure side of the system. One accumulator 122, a 198 cubic inch accumulator, is mounted in the diffuser inner body. This accumulator carries a nitrogen precharge of 700 psia, and acts primarily as an oil reservoir to supply oil during periods of high oil demand. The second accumulator 1402, a 60 cubic inch accumulator, is mounted near the pump outlet in the missile hydraulic section. This accumulator 1402 carries a 1,000 psia nitrogen precharge and acts primarily to smooth out high pressure peaks.

In the missile system, each wing 112, 113, 114, or 115 is individually controlled. Thus, for each wing an actuator 1382, 1383, 1384, or 1385 and a hydraulic control valve 1392, 1393, 1394 or 1395 is necessary to control the wing. The control valves 1392 through 1395 are balanced spindle linear flow type valves which control oil flow to the actuators 1382 through 1385. The position of the valve spindles is controlled by the previously mentioned torque motors 1378 through 1381, which are of the linear displacement, electromagnetic type. Thus, in the design range the output of the transfer valve-torque motor combinations, oil flow is linear with differential current applied to the torque motor coils.

The transfer valve-torque motor combinations control the flow of hydraulic oil to the wing actuators 1382 through 1385, which is the power dissipating portion of the hydraulic system 220. The actuators 1382 through 1385 consist of a hydraulic piston, mounted in an oil tight cylinder. Each end of the actuator is connected to one side of one of the transfer valves 1392 through 1395. A piston rod 1386, extends through one end of the actuators 1382 through 1385 and is connected to the wing bell crank 329 by bolt 1388.

All oil entering the transfer valves 1392 through 1395 is passed through a 10 micron filter. Two filters 1404 and 1406 are used, with each filtering the oil to two pairs of transfer valves 1392, 1393, and 1394, 1395, respectively.

Low pressure oil from the actuators 1382 through 1385 after passing through the transfer valves 1392 through 1395, is fed to the missile sump 1396. The sump 1396 is a cylindrical tank containing a nitrogen bladder which is precharged to 20 psia. The sump 1396 acts as a reservoir to store oil necessary for charging the high pressure accumulators 122 and 1402. It provides a pressure head to the inlet side of the hydraulic pump 1390; a cooling volume for the oil after passing through the hydraulic system; and oil to make up for system leakage. The sump 1396 is fitted with a low pressure oil gauge 1408 and a small charging line 1410. When the hydraulic system 220 is once bled and charged, it is maintained under pressure by the nitrogen precharge on the sump bladder 1412. Leakage from the system 220 can be made up by precharging the sump 1396 through the sump charging line 1410.

Prior to flight of the missile, the sump 1396 is charged to a pressure of approximately 100 psia. Immediately prior to firing, the hydraulic system sump pressure is checked, and the hydraulic system 220 is recharged if the sump pressure is 65 psia or lower at the time of firing.

The hydraulic turbine-pump assembly 1281 is the power source of the hydraulic system. The rotor of turbine 1282, illustrated in detail in FIG. 76, revolves at a very high speed, and consequently, has to be geared down through a reduction gear assembly 1400 to the lower pump operating speeds. A housing 1414 for the turbine assembly 1282 and the reduction gear assembly 1400 is made of an aluminum alloy casting. It is machined internally to provide a casing for the turbine assembly 1282 and a base and cover for the reduction gear assembly 1400. The forward end of the housing 1414 is secured to the outlet 1416 of the turbine air intake strut 774 by a plurality of bolts. The after end provides a base for the hydraulic pump 1390.

The turbine 1282 is a single stage, axial flow impulse type. A built-up turbine wheel 1418 is constructed of a disc of aluminum with 36 slots 1420 milled in the circumference. Steel blades 1422 are staked in these slots 1420. This wheel is carefully balanced and spin-tested to 50,000 rpm. The wheel 1418 is secured to a tapered section 1424 of a shaft 1426 by a lock ring 1428. This lock ring 1428 engages two milled slots in the face of the wheel 1418 and a flatened surface of the shaft 1426. Lock nuts 1430 hold the lock ring 1428 and wheel 1418 in position axially.

A nozzle plate 1432 for the trubine assembly 1282 is made similar to the trubine wheel 1418. Sixteen blades 1434, forming nozzles 1435, are machined in an aluminum disc 1436. This disc fits inside the housing 1414 and it is locked in place by a pin (not shown) and a stainless steel ring 1438 bolted to the housing 1414. A heavy duty, ball type, thrust bearing 1440 is fitted in an annular groove 1442 in the forward face of the nozzle plate 1432 to help support a throttle plate 1444.

The throttle plate 1444 is provided to control the volume of air to the turbine wheel 1418, which, in turn, controls the outlet pressure of the pump 1390 and the oil flow. Plate 1444 is a disc of aluminum. Slots 1446 are machined in the outer edge of the plate 1444 to coincide with the inlets of the nozzles 1435. The forward race 1448 of the thrust bearing 1440 is fitted into an annular groove 1450 in the after face of the throttle plate 1444.

An oilite bushing 1452, which is located between a shaft 1454 which is, in turn, bolted to the nozzle plate 1432 and a shell 1456 which is, in turn, bolted to the throttle plate 1444, provides a radial bearing surface. One thirty-second of a revolution of the throttle plate 1444 changes the opening of the throttle plate from a fully opened to a fully closed position. The position of the throttle plate 1444 is controlled by a pump control assembly 1458.

In order to protect the turbine 1418 from damage by foreign matter, stainless steel screens 1460 and 1462 are fitted to the air inlet and outlet of the turbine assembly 1282.

A bearing housing 1464, formed of cold rolled steel, is fitted into and secured to the outer housing 1414 by bolts 1466. It provides support for the turbine shaft 1426 by utilizing two rows of ball bearings 1468, separated by a carbon face oil-seal assembly 1470. The bearings 1468, oil-seal 1470, a spacer 1472 and a pinion gear 1474 are fitted over the shaft 1426 and held against a shoulder 1476 by lock nuts 1478. This assembly fits inside the bearing housing 1464 and is held in place by a snap ring 1480 bearing against the outer race of the after bearing 1468. An end cap 1482, bolted to the bearing housing 1464, holds the assembly in position.

A pinion gear 1474, keyed to the shaft 1426, meshes with three idler gears 1484. The idler gears 1484, fitted with needle bearings 1486, rotate on shafts 1488 which are held in drilled holes 1490 in the bearing housing 1464 by lock screws 1492. The idler gears 1484 mesh with and drive a bell gear 1494. The bell gear 1494 is made up of a ring gear 1496 fitted and pinned to a bell shaped body 1498. The body 1498 is machined with a stub shaft 1500 inside the bell and a male spline drive shaft 1502 on the outside thereof.

Two rows of ball bearings 1504 fit over the stub 1500 and inside the bearing housing 1464. These bearings 1504 are secured in position by snap rings 1506 and 1508 and an end cap 1510 bolted to the bearing housing 1464. By means of the 15 tooth pinion gear 1474, 39 tooth idler gears 1484, and a 93 tooth bell gear 1494, a speed reduction of 6.2 to 1 is obtained and the rotation is chagned from counterclockwise to clockwise.

An inductance type electrical pick-up 1512 is mounted in the outer housing 1414 near the bell gear 1494. This pick-up 1512 is used to determine the pump speed in missile testing and flight telemetering.

The hydraulic pump 1390, as best seen in FIGS. 77, 78, and 79, is a seven piston positive displacement type pump. Its rated capacity is 20 gal/min at 1,750 psi discharge pressure at 4,000 rpm.

A housing 1514 for the pump 1390 is constructed of two machined aluminum alloy sections 1516 and 1518. These sections 1516 and 1518 are secured together by bolts 1520 and are sealed with an O ring 1522. The forward casting 1516 is machined to house a stationary beryllium copper cam plate 1524. The cam plate 1524 is secured with a dowel pin (not shown) at an angle of 16.4.degree. from the perpendicular of the pump axis. A morganite bushing 1526, doweled to the casting 1516, is the forward bearing for the rotor 1548. An oil seal 1528 prevents hydraulic oil leakage from the pump 1390 to the reduction gear assembly 1400. A snap ring 1530 locks the oil seal 1528 and bushing 1526 in the casting 1516.

The after section 1518 of the housing 1514 is machined internally to receive a chrome steel port plate 1532, a spider washer 1534 and a large morganite bearing 1536. The port plate 1532 and morganite bearing 1536 are dowled in place and held in the after section 1518 with a snap ring 1538. Four bosses 1540, 1542, 1544 and 1546 are cast to the end of the after section 1518. The largest of these bosses, namely boss 1540, is the low pressure inlet to the pump 1390; while the next larger boss 1542 is the pump high pressure outlet connection. The two small bosses 1544 and 1546 are high and low pressure taps for use with the turbine control assembly 1458.

The rotor 1548 carries seven pistons 1550, only one of which is shown, in cylinders 1552, which are evenly spaced around a circle concentric with the rotor drive shaft 1554. These cylinders 1552 are fitted with silver carbon bushings 1556 in which the pistons 1550 move. The closed ends of the pistons 1550 are machined with a ball 1558 which is received in socketed slippers 1560 riding on cam plate 1524. A rotating auxiliary cam plate 1562 and springs 1564 mounted within the pistons 1550 hold the slippers 1560 against the cam plate 1524. The auxiliary cam plate 1562 is mounted on the rotor shaft 1554 by means of a spherical sleeve 1566, which is loaded by a spring 1568 to hold it aft against a flange 1570 on the slippers 1560. The springs 1564 are held in the rotor body 1548 and are prevented from rubbing on the port plate 1532 by spring retainers 1572 fitted in the end of the rotor cylinders 1552.

Small holes 1574 and 1576 drilled through the balls 1558 and the slippers 1560 provide forced lubrication to the bearing surface on the stationary cam plate 1524 during the piston pumping cycle. The port end of the rotor 1548 is ground and heavily plated with silver to provide a pressure seal against the port plate 1532. The port plate 1532 is a chrome steel disc, and it contains two kidney shaped ports 1578 and 1580 which are machined concentric with a center hole 1582. One of these ports 1578 serves as the pump inlet port, while the other port 1580 serves as the pump outlet port. The center hole 1582 in the port plate 1532 provides a passage to the pump inlet port 1540 to vent the internal pump leakage. The rotor 1548 is drilled with three passages 1584 to vent the cam plate end of the pump housing 1518 to the center hole 1582 of the port plate 1532.

When the pump 1390 is operating, the rotor 1548 rotates, carrying with it the seven pistons 1550, the auxiliary cam plate 1562, and slippers 1560. The slippers 1560 are held against the stationary cam plates 1524 by the piston springs 1564 and the auxiliary cam plates 1562. As the rotor 1548 rotates a given piston 1550 is moved axially in its cylinder bushing 1556 by the action of the stationary cam plates 1524. As the piston 1550 moves away from the port plate 1532, the open end of its cylinder 1552 is aligned with the kidney shaped inlet port 1578. When the piston 1550 reaches its extreme travel, it passes the plate area between the kidney ports 1578 and 1580. As the end of the cylinder 1552 aligns with the secondary kidney port 1580 (the outlet port), the piston 1550 is caused to move toward the port plate 1532 by the stationary cam plate 1524, and it pumps oil into the outlet port 1580. During this stroke, a small amount of oil is pumped through the small holes 1574 and 1576 in the balls 1558 and the slipper 1560, respectively, to lubricate the stationary cam plate 1524.

When the piston 1550 reaches the extreme end of its discharge stroke, the cylinder 1552 is covered by the plate 1532 between the kidney ports 1580 and 1578. As the rotor 1548 rotates further, the cylinder 1552 again aligns with the inlet port 1578 and the stationary cam plate 1524 again and causes the piston 1550 to move away from the port plate 1523, starting a new pumping cycle.

The pump control assembly 1458, as best seen in FIGS. 15, 80, and 81, regulates the discharge pressure of the pump 1390 by actuating the throttle plate 1444, thereby controlling the volume of air to the turbine wheel. The assembly 1458 is doweled and bolted to the turbine housing 1414 in such a manner that an arm 1586 on the control assembly 1458 actuates an arm (not shown) on the throttle plate 1444 by means of a rod 1588.

Oil from the pump discharge high pressure boss or top 1546 enters the assembly 1458 through a small porex filter 1590 and acts on the top of a small spring loaded servo piston 1592. The servo piston 1592 is held in the closed position by a compression spring 1594 acting on its lower end. This spring 1594 is used to adjust the pump discharge pressure. As the hydraulic pressure on the top of the servo piston 1592 overcomes the pressure of the spring 1594, the servo piston 1592 is displaced and it uncovers a port 1596 in the servo piston sleeve 1598. This allows oil to flow through passage 1600 in the body 1602 of assembly 1458 to a spring loaded control piston 1604.

A hole 1606 drilled in the control piston 1604 bleeds oil from the top of the control piston to a control linkage area, and through a damping bleed 1608 to the pump inlet 1540 by way of the boss 1544. Thus, the pressure applied to the control piston 1604 is determined by the bleed circuit formed by the opening of the servo piston 1592 and the hole 1606 in the control piston 1604. Pressure applied to the top of the control piston 1604 moves the control piston against spring pressure and the throttle plate 1444.

When no pressure is applied by the servo piston 1592, a control piston spring 1612 moves the control piston 1604 and throttle linkage 1610 to open the throttle plate 1444 and pass more air to the turbine wheel 1418. Motion of the control piston 1604 is applied to the throttle plate 1444 through a crank arm 1614. Crank arm 1614 is hinged to the control piston 1604 and it is slidably received on a crankshaft 1616. Movement of the hinged end of the crank arm 1614 rotates the crankshaft 1616, which, in turn, rotates the throttle plate control arm 1586. The throttle plate 1444 is secured to the control arm 1586 by a rod 1588, the ends of which are provided with ball and socket joints 1618. When the control piston 1604 is moved its full stroke, the throttle plate 1444 revolves from a full open to a full closed position.

The purpose of the hydraulic sump 1396, FIG. 82, is threefold, namely, to provide oil for accumulator storage and system leakage; to provide pressurized oil to the inlet of pump 1390 to prevent cavitation thereof; and to provide a reservoir for cooling oil.

The cylindrical shell 1620 is sealed at one end by an end plate 1622. The end plate 1622 is drilled and tapped for a system supply line 1624, a system drain line 1626, three return lines 1628, 1630, andd 1632, and an oil precharge line 1410.

The opposite end of the cylinder 1620 is sealed by bladder support ring 1634 which is affixed to shell 1620 of the sump 1396. The bladder cover plate 1636 is held in place by a threaded ring 1638 screwed into the bladder support 1634. The cover plate 1636 is drilled and tapped to receive a high pressure air valve 1640, which is used to precharge the bladder 1412 with nitrogen pressure. The high pressure air valve 1640 is a double seal valve. One seal of the valve is a spring loaded stem and disc, which is similar to a conventional tire type air valve. The other seal of the valve is a metal to metal seal which is seated manually with a wrench inside the sump 1396.

A baffle plate 1642, containing a plurality of holes 1644, is located near end plate 1622. This baffle plate 1642 serves as a bladder stop; prevents the bladder 1412 from expanding the full length of the sump 1396; and maintains a reservoir of oil for cooling.

The bladder 1412, contained in the sump 1396, is made of synthetic rubber and it is formed in the shape of a bellows with one open end. The neck 1646 of the open end of the bladder 1412 is seated in the bladder support 1634, with the bladder cover 1636 sealing the bladder. The opposite end of the bladder 1412 is free. The bladder 1412 will expand longitudinally when pressurized. The expansion and collapsing of the bladder is controlled by the variation in bladder wall thickness. The bladder 1412 is limited to a working volume by the baffle plate 1642.

To charge the sump 1396 for use, the bladder 1412 is charged first with nitrogen to a pressure of 20 psia with the oil side open to atmosphere by a vent 1648. This will cause the bladder 1412 to expand to the baffle plate 1642. The system is then charged with oil through the oil precharge line 1410 to 100 psia. This will force oil into the sump 1396 and compress the bladder 1412 until the nitrogen pressure is equal to the oil charge pressure. When the system pressure is at its maximum pressure of 1750 psia, oil will be removed from the sump 1396 and stored in the two accumulators 122 and 1402. This allows a reserve storage of oil to compensate for system leakage.

In the hydraulic system 220, two high pressure accumulators are used. These accumulators 122 and 1402 are used to supply oil for high energy storage to meet peak flow demands, and to prevent pressure surges due to rapid change in oil flow demand and pump output pressure fluctuations.

Accumulator 122, is located in the inner body 120 of missile 100 as best seen in FIG. 7. This accumulator has a nitrogen precharge pressure of 700 psia on it, and it is used primarily for delivery of oil at high flow demands, especially at high altitudes. The second high pressure accumulator 1402 is located in compartment 220 just aft of bulkhead 308. This accumulator 1402 carries a nitrogen precharge pressure of 1,000 psia. Due to its proximity to the hydraulic pump outlet 1542, this accumulator 1402 acts primarily to smooth the pressure fluctuations of the hydraulic system.

The accumulators 122 and 1402 are identical in construction except for the dimensions of the shell 1650, bladder 1652 and bladder cover plate 1654. Each accumulator 122 or 1402 is comprised of a spherical shell 1650 to which there is affixed a fitting 1656 for connection with an oil line 1658. A perforated baffle plate 1660 is press fitted into the base of the oil connection fitting 1656 to support plug 1662 which is attached to the bladder 1652 when no oil is in the accumulator 122 or 1402. Also affixed to the shell 1650 is a boss 1664 containing the bladder seat 1666. This boss 1664 is threaded to receive a bladder cover securing ring 1668. The bladder 1652 is a synthetic rubber sphere with a greater wall thickness at the precharge side and a smaller wall thickness at the oil line side thereof. This difference in wall thickness controls the collapsing action of the bladder when the accumulator is charged with oil.

A series of small buttons 1670 are molded on the lower half of the bladder 1652, on the oil side, to prevent the bladder 1652 from sealing against the shell 1650 and trapping pockets of oil. A disc 1662 is secured to the end of the bladder 1652 at the point on the bladder which covers the oil connection. This disc 1662 prevents rupture of the bladder 1652 through the oil line 1658 when the bladder is charged with nitrogen. The disc 1662, which is secured to the bladder 1652, will move away from the plate 1660 when the hydraulic charge overcomes the bladder nitrogen charge, thus permitting the free flow of oil through the plate 1660 to the accumulator.

The neck 1672 of the bladder 1652 is held against the seat 1666, and it is sealed by the bladder cover 1654. The bladder cover 1654 is drilled and tapped to receive the previously mentioned high pressure air valve 1674. The cover 1654 contains a small orifice 1676 which restricts the rate of flow of nitrogen from the air valve chamber 1678 to the bladder 1652, thus preventing high nitrogen flows from the bladder 1652 in case of rupture of the air valve 1674.

The high pressure air valve 1674 is a double seal valve, one seal being a spring loaded stem and disc similar to a conventional tire type air valve, with the second seal a metal to metal seal which is seated manually with a wrench. The bladder cover 1654 is secured in place by the steel securing ring 1668 which screws into the boss 1664 on the shell 1650.

As shown in FIGS. 15 and 83, the control transfer valves 1392, 1393, 1394 and 1395, serve as the controlling units for flow of oil to the wing actuators 1382, 1383, 1384 and 1385, respectively. These valves control the velocity and position of the wings 112, 113, 114, and 115. The valves are of a linear flow spindle type. A valve spindle 1680 in each of the valves 1392, 1393, 1394 and 1395, is positioned axially by the torque motor 1378, 1379, 1380 or 1381, respectively. These motors are rigidly mounted on the valve assembly. The torque motors are controlled by the DC current from the servo amplifier in the guidance control system 214.

The transfer valve assemblies are all alike. Therefore, only one transfer valve assembly, such as 1392 illustrated in FIG. 83, will be described. The transfer valve 1392 comprises a body 1682 containing a sleeve 1684, a spindle 1680, end sealing plugs 1686 and 1687, a spindle seal 1688, damping chambers 1690 and 1692, and bleeds 1742 and 1744 which must be withdrawn to bleed air, if desired. The body casting 1682 forms the outer shell for the entire valve assembly 1392. A large hole 1696 is bored into the body casting. The sleeve 1684 and spindle 1680 are fitted in this hole 1696. Also drilled in the body casting 1682 are four hydraulic ports: an inlet port 1698, a drain port 1700, and two wing actuator ports 1702 and 1704. The hydraulic system pressure of 1750 psia is applied to the valve 1392 through the inlet (high pressure) port 1698. One of the lines 1628, 1630, or 1632, leading to the missile sump 1396, is attached to the port 1700. The remaining two ports 1702 and 1704 are connected to opposite ends of one of the wing actuators 1382, 1383, 1384 or 1385, through tubing and flexible connectors. Other holes drilled in the casting serve as bleeding ports and damping chamber.

The sleeve 1684 is comprised of two end sections 1706 and 1708, and a center section 1710. Metering ports 1712 and 1714, and inlet ports 1716 are contained in the center section 1710. The length of this section 1710 extends from one outside edge to the other outside edge of spindle lands 1718 and 1719.

The metering ports 1712 and 1714 are two diametrically opposite milled slots cut into the ends of the center section 1710 of the sleeve 1684. Four round inlet ports 1716, spaced equally around the circumference are drilled at the midpoint of the center section 1710 and connect with the high pressure inlet port 1698 in the casting 1682. The end sections 1706 and 1708, as well as the center section 1710 of the sleeve 1684 are secured together and fitted into the body casting 1682. These sections are so oriented as to allow even distribution of oil flow to all the ports.

The spindle 1680 is the only moving part in the transfer valve 1392. The spindle 1680 is machined and ground so that the two metering lands 1718 and 1719, and the end lands 1720 and 1721 have a very smooth finish. The spindle 1680 is hand lapped in to the sleeve 1684. A hole (not shown) is drilled axially through the center of the spindle 1680 to receive a positioning wire 1722. The positioning wire 1722 is silver soldered into place before the finish grind operation.

The positioning wire 1722 serves as the mechanical connecting link between the transfer valve spindle 1680 and the armature of the torque motor. It is used to transmit the movement of the armature to the spindle 1680.

Two damping chambers 1690 and 1692, one at each end of the valve sleeve 1684, are formed by the valve spindle outside edges and the sealing plugs 1686 and 1687 which are fitted into the body casting 1682 at each end of the main body casting sleeve hole 1696. The two damping chambers 1690 and 1692 are filled with hydraulic oil and vented to drain pressure through a feeding orifice 1724. The damping chambers 1690 and 1692 are connected together by a passage 1726, which is drilled in the main body casting 1682. This passage 1726 contains a damping orifice 1728, which aids in damping the valve spindle oscillations and permits equalization of the pressure on each end of the spindle 1680 when it is moved.

The sealing plugs 1686 and 1687, used to close the damping chambers 1690 and 1692, respectively, are screwed into the main body casting 1682 and are sealed by O rings 1730. The sealing plug 1687, at the end of the spindle 1680 which carries the piano wire linkage 1722, carries a positioning wire seal 1688. This seal 1688 includes a small washer 1732, which has a center hole of the same diameter as the positioning wire 1722. The washer 1732 fits into a hole 1734 in the sealing plug 1687 and is held in place by a self-aligning spacer 1736, and a retainer 1738 secured to the plug 1687 by two machine screws 1740. The hole 1734 in the sealing plug 1687 is of such a size that the washer 1732 can be moved to obtain minimum eccentricity and it is then locked into place.

The plug 1686, at the opposite damping chamber 1690, is fitted with the bleed screw 1742 at its center to bleed air from the damping chamber 1690. A second bleed point is provided by a bleed screw 1744 at the end of passage 1726. This bleed screw 1744 is sealed by an O ring 1746, and it is drilled to permit removal of air from the transfer valve 1392 by merely loosening the screw 1744.

The feeding orifice 1724 connects the damping chamber 1692, through which the positioning wire 1722 runs, to the drain port 1700. This orifice 1724 permits oil flow to and from the damping chamber 1692 to compensate for volumetric change between the two damping chambers 1690 and 1692 due to the positioning wire 1722. This bleed 1724 also allows the pressure in the damping chambers 1690 and 1692 to equalize with the drain pressure.

The torque motors 1378, 1379, 1380 and 1381, FIG. 84, which are the actuating devices for the transfer valves 1392, 1393, 1394 and 1395, respectively, convert an electrical signal from the guidance control system 214 into a linear motion proportional in direction and amount to the polarity and amplitude of the impressed electrical signal. This linear motion is transmitted to a transfer valve such as valve 1392, by the wire 1722. The torque motors used in the missile 100 are linear displacement, spring balanced devices whose output motion is directly proportional to the impressed electrical signal.

The typical type torque motor 1378 is comprised of an armature 1748 supported by a torsion spring 1750 between two sets of magnetic poles 1752 and 1754. Coils 1756 and 1758 surround the armature 1748. The two sets of magnetic poles 1752 and 1754 that the armature 1748 works against are formed by two sets of heavy, highly magnetized, flat permanent magnets. The coils 1756 and 1758 surrounding the armature 1748 are so connected that when current flows in one direction, the armature 1748 is deflected in one direction, and when current flows in the opposite direction, it is deflected in the other direction.

The magnetic circuit, coils 1756 and 1758, as well as the torsional spring 1750 are so constructed that the motion of the driving end of the armature is directly proportional to differential current applied to the coils 1756 and 1758. The armature is balanced by a weight 1760, mounted on the end of then armature 1748 opposite the transfer valve connection 1722. This balance reduces motion of the end of the armature due to a high frequency vibration.

The operational sequence of the torque motor 1378 and transfer valve 1392 will now be discussed. A DC differential current from the guidance control system 214 is applied to the torque motor coils 1756 and 1758, causing the torque motor armature 1748 to move. This motion, transmitted to the transfer valve spindle 1680 by the connecting wire 1722, causes the spindle 1680 to move. When the spindle 1680 moves, land 1719 uncovers one set of metering ports 1714 to the drain port 1700, while the other land 1718 uncovers the other set of ports 1712 to the high pressure oil contained between the spindle lands 1718 and 1719 from the inlet port 1698.

Thus, oil flows from the valve inlet port 1698, across the valve spindle 1680 between the lands 1718 and 1719 to the metering ports 1712 connected to high pressure, through the metering ports 1716 and thence to outlet port 1702 to one side of a wing actuator, such as actuator 1382. As the actuator 1382 moves, oil from its opposite side is forced to the transfer valve 1392, through the other set of metering ports 1714 to the valve drain port 1700 and the missile sump 1396.

The wing actuators 1382, 1383, 1384, and 1385 are hydraulic-mechanical devices which serve to move the wings 112, 113, 114, and 115, respectively, and hold them in any position from 20.degree. clockwise to 20.degree. counterclockwise from the longitudinal axis of the missile 100. Hydraulic flow of the oil to and from the actuators 1382 through 1385 is controlled by the transfer valves 1392 through 1395, respectively.

As illustrated in FIGS. 20, 21, and 85, an actuator assembly, such as actuator 1382, is composed of a body 1764, a cylinder cover 1766, a sealing nut 1768, a piston assembly 1770, and sealing O rings 1772 and 1774. The piston 1770 of the actuator 1382 has two grooves 1776 machined on its periphery for piston rings 1778.

A mechanical stop 1780 is machined in the closed end of the body casting 1764 for the piston 1770. A piston rod hole 1782 is machined in the center of the cylinder cover 1766. This hole 1782 is grooved to permit the O ring 1774 to be fitted between the piston rod 1386 and the cylinder cover 1766 to seal against oil leakage past the piston rod. Oil leakage between the cylinder cover 1766 and the body casting 1764 is also sealed against by the O ring 1772. The cylinder cover 1766 acts as one of the mechanical stops for the piston 1770.

A cylinder nut 1768 holds the cylinder cover 1766 and the body 1764 together. The cylinder nut 1768 fits over the cylinder cover 1766 and locks the cover to the main body 1764. The nut 1768 also permits the cover 1766 to be moved radially to insure that the piston rod 1386 is aligned.

The piston rod 1386 is threaded at the end and it has a keyway (not shown) cut in it for a castellated nut that prevents rotation of the piston 1770. The threaded end of the piston rod 1386 is screwed into a threaded eye 1784 and locked in place by means of a nut 1786 and a lock washer 1788.

The piston rod 1386 is adjusted to give equal throw to the wings 112, 113, 114, and 115 by screwing it in or out of the threaded eye 1784, and locking it in place by the lock nut 1786 and lock washer 1788. Two bosses 1790 and 1792, one on the body casting 1764 and one on the hydraulic cover casting 1766, are threaded to receive tubing 1794 from the control transfer valves 1392 through 1395.

In operation, hydraulic oil is applied to one side of piston 1770 causing it to move, thereby moving the wing 112, 113, 114, or 115, through the bell crank 329, and displacing the oil on the opposite side of the piston 1770.

In order to maintain the aerodynamic alignment of the missile 100 during the boost phase when the servo system is inactivated and sufficient hydraulic pressure is not available in the wing actuators 1382 through 1385 to hold the wings 112, 113, 114, and 115 at the zero position, it is necessary to pin the wings in their zero position. This is done by the previously mentioned four hydraulically operated wing locks 140, one for each wing mounted on the wing lock struts 142.

As illustrated in FIGS. 25, 86, and 87, each wing lock body 1796 contains an internal cylinder 1798, which is externally threaded to receive an end cover 1800, a boss 1802 threaded for a hydraulic oil line 395, and brackets 416 for the piston latch 409. A piston 1804 and a piston spring 1806 fits into the wing lock body 1796. The wing lock piston 1804 is formed integrally with a locking pin 1808, a land 1810, and the previously mentioned guide and latching pin or rod 397.

One end 399 of the locking pin 1808 is machined with a tapered flat 1812 on each side to form a wedge which engages a recess 401 in the wing rib 311, as previously discussed. The land 1810 is spaced from the other end of the piston 1804. An O ring 1814 is seated in a groove 1816 adjacent this land 1810 to form a seal between the land and the wall of the cylinder 1798. The latching pin 397 of the piston 1804 has the previously mentioned notch 413 provided therein and a keyway 1818. The keyway 1818 cooperates with a pin (not shown) which is seated in the cover 1800, to prevent the piston 1804 from turning.

The wing lock spring 1806 is fitted over the guide and latching pin 397 and is seated within the piston land 1810 where it is held in position by the wing lock end cover 1800. The cover 1800 is internally threaded to mate with the threads on the body 1796. The piston guide and latching pin 397 passes through a hole 1820 provided in the cover 1800. The piston latch 409 is fabricated from steel plate. This latch is claw shaped, with a lever arm and the previously mentioned pawl 411. The latch 409 is provided with notch 407 to receive an unlatching tool. The pawl 411 is tapered to a nose which seats in the latching pin notch 413. The latch 409 is spring loaded by leaf spring 417 to bear against the guide and latching pin 397, as best seen in FIG. 25.

After the wing lock 140 is assembled and secured to the wing lock beam 142, the locking pin 1808 is forced into the wing lock body 1796, compressing the wing lock spring 1806. This extends the guide and latching pin 397 to the position where the pawl 411 engages the notch 413 on the guide and latching pin 397. The pawl 411 of latch 409 holds the piston 1804 against the spring tension until the latch 409 is manually retracted against its spring 417 by inserting an unlatching tool through the beam 142, as previously mentioned. When the latch 409 is retracted, the wing lock spring 1806 will move the piston 1804 to extend the locking pin 1808 into the wing, such as 112, 113, 114, or 115.

To retract the locking pin 1808 upon separation of the missile 100 from the booster 500, a solenoid valve 1820, illustrated in FIGS. 15 and 91, is activated by a separation circuit which applies hydraulic system pressure to the top of the wing lock piston 1804. The hydraulic pressure acts on the piston land 1810 to overcome the tension of spring 1806 and moves piston 1804 so that the locking pin 1808 is retracted from its wing while the latch 409 engages the guide and latching pin 397. The wing lock 140 is then held in an unlocked position by the latch 409 in case of loss of power to the solenoid valve 1820.

The solenoid valve 1820 is a three way ball type valve operated by solenoid 1822. This solenoid is energized at separation of the missile 100 from the booster 500 by the separation circuit and it applies hydraulic system pressure to the four wing locks 140 and a pressure switch 1824, as indicated in FIG. 15.

The valve body 1826 is drilled and tapped for three oil connections at bosses 1828, 1830, and 1832. The bottom boss 1828 is used in conjunction with the hydraulic system pressure. Boss 1830 is connected to the supply line 395 to the wing locks 140 and the pressure switch 1824, while boss 1832 is connected to the sump return line 1850.

A split sleeve is machined to form a valve cage 1834. This cage 1834 is seated in the valve body 1826 and sealed with O rings 1836 and 1838. The bottom piece 1840 of the cage 1834 contains an inlet port 1842 and a seat 1844, while a top piece 1846 thereof contains four ports 1848 which communicate with the wing lock line 395 and the sump return line 1850. The top piece 1846 also contains return ports 1852 and a seat 1854, and it serves as a guide for a push pin 1856. A valve ball 1858 is contained within the cage 1834. The valve cage 1834 is held in place by a retainer 1860 machined from stainless steel which is screwed into the valve body 1826. A pin 1862, secured in the body casting 1826, engages a notch 1866 in a flange 1868 on retainer 1860, locking it in position. The retainer 1860 also serves as a guide for a solenoid armature 1864.

The solenoid armature 1864 is a disc type armature, and it contains ports 1870 for oil passage through it to the sump 1396. A spring 1872 is seated within the armature 1864 to hold the armature, push pin 1856, and ball 1858 in a normally closed position. The spring 1872 is positioned by a chrome button 1874 secured to a diaphragm 1876. The diaphragm 1876 is a thin disc which is seated in the valve body 1826 to form an oil tight seal between the valve and a solenoid coil 1878. Solenoid housing 1880 is seated on the diaphragm 1876 and it is secured by four bolts 1882 to the valve body 1826. The solenoid coil windings 1878 and a core 1884 are contained in the solenoid housing 1880 and energized through two wire leads 1886 and 1888.

When the solenoid 1822 is not energized, the compression of spring 1872 will hold the ball 1858 against the inlet seat 1844 and permit oil flow from the wing locks 140 and pressure switch 1824 to the sump 1396. This is the position during the boost phase of the missile 100. On separation of the missile 100 from the booster 500, the solenoid coil 1878 is energized, forming a magnetic field within the core 1884 which lifts the armature 1864 against the spring 1872. This permits the hydraulic system pressure to raise the ball 1858 and allow oil to flow to the wing locks 140 and pressure switch 1824. This hydraulic pressure will hold the valve ball 1858 against the outlet seat 1856, sealing the sump return ports 1852.

Three cone type, spring loaded, check valves 1890, 1892, and 1894, as illustrated in FIG. 88, are used in the hydraulic system 220. The function of these check valves is to insure that oil flows in one direction only at three critical locations in the hydraulic system 220. These check valves are conventional fittings.

Check valve 1890 is installed in the sump return line 1850 from the solenoid valve 1820 and it functions to prevent premature action of the wing locks due to the pressure in the sump 1396. This valve 1890 also serves to prevent draining of oil from the wing locks 140 to the sump 1396 during periods of shut-down of the hydraulic system 220 when no pressure is on the sump 1396.

Check valve 1892 is installed between the high pressure lines of the hydraulic system 220 and the external high pressure line 1896 to the solenoid valve 1820. This check valve 1892 prevents the entire hydraulic system 220 from being pressurized when external power is used to operate the wing locks 140.

Check valve 1894 is installed in the pump discharge line 1898 to prevent reverse motoring of the pump 1390 during tests when an external hydraulic power supply is used. When the hydraulic system 220 is operated with internal power and load requirements are such that pump discharge pressure is less than accumulator pressure, this check valve 1894 will prevent the accumulators 122 and 1402 from discharging through the pump 1390 to the sump 1396.

A cone type disc 1900 of the check valves 1890, 1892, and 1894 is drilled and tapped to receive a bleed 1902, FIG. 88. This bleed 1902 is included to permit air entrapped in the pump 1390 and throttle mechanism 1458 to pass to the high pressure portion of the hydraulic system 220 where it can be removed by bleeding. This bleed 1902 passes sufficient oil to the pump 1390 to cause the pump to operate at low speed when the hydraulic system 220 is operated on external hydraulic power. This low speed is not sufficient to cause damage to the pump 1390.

Four quick disconnect, self-sealing couplings 1904, 1906, 1908, and 1910 are used in the hydraulic system piping to provide external services, including external application of hydraulic pressure during factory and field testing; to provide hydraulic drain during factory and field testing; to operate the wing locks 140 during subsystem testing; and to charge the sump 1396.

The four quick disconnect couplings 1904 through 1910 are standard commercial items. Therefore, a detail description of the couplings is not provided.

Two large filters 1404 and 1406, as illustrated in FIG. 89 each incorporate a cartridge 1912 and, are used in the hydraulic system 220 to assure a dirt-free hydraulic oil supply to the transfer valves 1392 through 1395 and the wing actuators 1382 through 1385.

One filter 1404 is used to filter all oil flow to the transfer valves 1392 and 1393 which control wings 112 and 113, respectively. The other filter 1406 is used to filter all flow to the wing locks 140 and to the transfer valves 1394 and 1395 which control wings 114 and 115, respectively. The body 1914 of the filters 1404 and 1406 contains a spring loaded relief by-pass valve 1916, located between an inlet port 1918 and an outlet port 1920.

In addition to the two large filters 1404 and 1406, two smaller filters 1922 and 1924, FIG. 90, are used in the external lines 1896 and 1410 to the hydraulic system 220. One filter 1924 is used in the sump charge line 1410 and the other filter 1922 in the wing lock external supply line 1896. A relief valve disc 1926, loaded by spring 1930 supports filter element 1928.

The operation of the hydraulic system 220 in connection with aerial flight of the missile 100 will now be described. As soon as the missile-booster combination begins to accelerate, the air scoop 772 provides air through channel 1280 in the intake strut 774 for the hydraulic pump 1281 so that the turbine 1418 will begin rotation. The throttle control piston 1604, since no hydraulic pressure exists, will be in full throttle position. Consequently, the turbine 1418 will accelerate rapidly. Since the pump 1390 and turbine 1418 are connected directly through a gear reduction assembly 1400, the pump 1390 will immmediately begin pumping oil.

When the pressure of the hydraulic system 220 has reached 1,000 psia, any further pump stroking will not be acting against a non-yielding system, since the oil pressure is now greater than the accumulator pre-charged nitrogen pressure. Consequently, the accumulators 122 and 1402 will consume oil, and as a result of increased oil storage, the nitrogen will be compressed. The accumulator pressure will, therefore, rise as the nitrogen volume decreases. The net yield will be system pressure rapidly rising to nominal value, while the pump 1390 is storing oil in the accumulators 122 and 1402 at the same pressure.

The oil volume stored in the accumulators 122 and 1402 must, of course, come from the pressurized sump 1396. During the system pressurizing period, the sump 1396 will have discharged oil, and, consequently, the sump pressure will decrease. The sump 1396 will contain enough oil to keep the intake pressure to the pump 1390 above 20 psia and thus will prevent cavitation.

When the pressure of the hydraulic system 220 has reached a level of 1,750 psi, the throttle control piston 1604 will close, thus decelerating the turbine 1418. If all action described above takes place before the boost phase of the missile 100 is completed, the speed of the turbine 1418 may actually decrease quite rapidly. If, however, the boost phase of the missile 100 is completed, the hydraulic system 220 will demand some oil flow which will not allow the turbine 1418 to slow down as rapidly, or if demand is great enough, may even call for more speed from the pump 1390. However, on the basis of flight data and calculations, the hydraulic pressure should actually be 1,750 psi some time before the boost phase of the missile 100 is completed.

At the time of separation of the missile 100 from the booster 500, tail end separation switches will close, thus energizing the solenoid valve 1820 in the wing lock sub-system 1398. The valve 1820 will supply system pressure to the wing locks 140 and the wing locks will retract, leaving the wings 112, 113, 114, and 115 free to rotate.

The missile guidance system 214 will also energize at separation of the missile 100 from the booster 500. Undoubtedly response to a relatively large roll error will be required at separation of the missile 100 from the booster 500 due to missile roll misalignment encountered during boost. The pressure of the hydraulic system 220 will decrease because of this demand, and, as a consequence, the pump throttle control piston 1604 will again permit more air to flow to the turbine 1418. The turbine 1418 will accelerate to keep up with the demand, and, in the interim, the accumulators 122 and 1402 will supply oil to the system. In the event of a decrease in flow demand, the turbine 1418 will decelerate to the speed corresponding to the new flow requirements and, during this deceleration period, the accumulators 122 and 1402 will store the excess oil.

Assuming initial capture of the missile 100, the mid-course guidance phase will probably be the quietest period of operation for the hydraulic system 220 during flight of the missile. The operation of the hydraulic system 220 will probably be a somewhat milder form of the capture phase characteristics. All flow to the wing actuators 1382, 1383, 1384, and 1385 will pass through the filter-transfer valves circuit and then into the sump 1396 for re-cycling through the pump 1390. Some cooling will take place in the sump 1396, although this will be a direct function of quantity of oil in the sump 1396 and thus of the pressure of the hydraulic system 220.

The terminal guidance phase of the missile 100 will probably be the most severe case for the hydraulic system 220. This is a result of the following conditions:

1. Noise level in the homing guidance sub-system is expected to be high, and consequently, jitter in the wings 112, 113, 114, and 115, will probably result. This will call for large demands of oil flow from the valves 1392 through 1395;

2. Air pressure will be at its lowest value, and, consquently, available power from the pump 1390 will be at its lowest value. Response time of the system will undoubtedly be greater; and

3. Oil temperature will be at its highest value.

The missile constituting the subject matter of the instant invention may incorporate any one of the several well-known guidance and/or homing systems.

The booster 500, illustrated in FIGS. 44 through 61, utilized for propelling the missile 100 to a nominal separation velocity of approximately 2000 ft/sec is comprised of a large, solid propellant rocket which provides approximately 370,000 lb/sec of impulse for a period of approximately 4 seconds. This booster 500 supports the missile 100 before launching from a zero length trainable launcher 1000, FIG. 61 (or a dual launcher for shipboard use). The weight of the booster 500 less its two pairs of fins 257 and 259, and 256 and 258, is estimated at approximately 2,850 pounds. The pairs of fins 257 and 259, and 256 and 258 are attached to the booster 500 for longitudinal stability of the missile-booster combination.

The case 501 for the booster 500 is a cylinder. At one end 3999 of this cylindrical case 501, there is attached a convergent-divergent nozzle 4000 by means of threaded bolts 4002, FIG. 60. The end 3999 of case 501 supports aft launching and handling shoes 4005 and 4007, respectively. A gasket 4004 is positioned between a flange 4006 of the nozzle 4000 and the inner surface of the thickened end 3999 of the booster case 501 to prevent gas leaks.

The nozzle 4000 is manufactured in two sections 4038 and 4040 which are welded together at 4042 just aft of the throat 4044 of the nozzle, as best seen in FIG. 60. Heat resistant inserts (not shown) can be provided in the nozzle 4000, or other means can be provided for cooling the nozzle 4000.

The exterior of the nozzle 4000 is encased in a shroud 4046 to reduce air drag. This shroud 4046 is attached to a rearwardly extending annular flange portion 4047 of the nozzle 4000 by means of bolts 4049, with an annular ring 4051 being secured between the shroud 4046 and the flange portion 4047 of the nozzle 4000 to reinforce the structure.

A nozzle closure disc 4003 is inserted in the nozzle exit 4052 to seal the nozzle 4000 in order to prevent loss of propellant solvents and to prevent entrance of moisture into the propellant grain 4184. It is manufactured of a light metal sheet and is generally inserted in the exit 4052 of the nozzle 4000, although it can be made smaller and inserted near the throat 4044 of the nozzle. When the propellant grain 4184 of the booster 500 is ignited, this disc 4003 is blown out of the exit 4052 of the nozzle 4000 after aiding in the initial build up of the pressure in the booster.

Two pairs of U-shaped annular channels 4060, mounted back-to-back (web-to-web), in spaced relationship, are utilized for securing the fin spars 4062 and 4064 from each fin 256, 257, 258 and 259 to the booster 500. Annular doublers 4066 and 4068, together with a cylindrical member 4070, are utilized to reinforce the channels 4060. The fin spars 4062 and 4064 are secured between the spaced pairs of channels 4060 by means of bolts 4072. A flange portion 4074 of each channel 4060 is secured to the shroud 4046 by welding or the like. The shroud 4046 is reinforced in the vicinity of each fin spar 4062 or 4064 by straps 4078 and 4080, respectively, which gird the shroud. Stringers 4083 are located between the inner channels 4060 to reinforce the shroud 4046, while an annular channel 4082 is utilized to reinforce the downstream end of the shtoud 4046.

At the other end 4010 of the booster case 501, a forward head 4012 (closure) is bolted thereto by means of bolts 4014, as best seen in FIGS. 46 and 47. A gasket 4016 is utilized to seal the shoulder portion 4018 of the forward head 4012 to the forward end 4010 of the booster case 501. This forward head 4012 supports an igniter 4020, an igniter arming device or mechanism 4022, a grain immobilizer 4024, a resonance rod assembly 4026, a plurality of clamping ring guards 4028, forward pairs of launching and handling shoes 4030 and 4032, respectively, as well as the booster thrust ring 4034 (formed integral with the forward head 4012) and clamping ring 4036, FIGS. 46 and 47.

As shown in FIGS. 44, 46, and 47, the missile 100 and the booster 500 are held together by means of the clamping ring 4036 whose flange portions 4091 and 4093 mate with machined surfaces 4092 and 4094 on the missile aft ring 4096 and booster thrust ring 4034, respectively. The missile aft ring 4096 is secured to the cylindrical section 284 of the missile by means of rivets 4097.

The clamping ring 4036 is so designed that, when it is unlocked by a striker arm 4098 on the launcher 1000, it springs outwardly into the six ring guards 4028 on the booster 500, thus unlocking the booster and missile 100. During the boost phase of aerial flight of the missile 100, the missile-booster combination is held together only by the thrust of the booster 500. The thrust is transmitted from the booster 500 to the missile 100 by mating machined surfaces 4099 and 4101 on the missile aft ring 4096 and the thrust ring 4034, respectively, as best seen in FIG. 47.

As illustrated in FIG. 46, the clamping ring 4036 is constructed of two semi-circular sections 4103 and 4105. One end of each section 4103 and 4105 of the clamping ring 4036 is secured to a bracket element 4107 or 4109, respectively, by means of bolts 4111. These bracket elements 4107 and 4109, in turn, cooperate with an adjustment screw 4113 to provide a turnbuckle 4114. Bifurcated bracket elements 4112 and 4115 are secured to the opposite ends of sections 4103 and 4105, respectively, of the clamping ring 4036 by bolts 4108, as shown in FIG. 46, 48, and 49.

A scissors hinge 4117, having its two arms 4119 and 4121 pivotally connected together by a pin 4123, is pivotally mounted to the bifurcated bracket elements 4112 and 4115 by pins 4110 and 4118, respectively. This scissors hinge 4117 permits concentric opening of the clamping ring 4036 so that the latter will unlock the missile aft ring 4096 and booster thrust ring 4034.

A bifurcated yoke 4116, as shown in FIGS. 48, 49, and 50, keeps the arms 4119 and 4121 of the scissors hinge 4117 together so that the sections 4103 and 4105 of the clamping ring 4036 are held in operative position. A shear pin 4131 is utilized to hold the yoke 4116 and arms 4119 and 4121 of the hinge 4117 together so as to prevent premature release of the sections 4103 and 4105 of the clamping ring 4036. When the striker arm 4098 releases the yoke 4116 by shearing the pin 4131, the two arms 4119 and 4121 of the scissors hinge 4117, are forced apart by leaf springs 4125, FIGS. 56, 57 and 58. These springs 4125 are mounted on the inner surface of the clamping ring 4036 and they force the clamping ring sections 4103 and 4105 outwardly to release the clamping ring as indicated above.

A bracket 4601, FIGS. 49 and 50, is secured to the booster casing 501 by suitable means, such as screws 4603 or the like. Two vertically extending and spaced eyelugs 4602 are provided as an integral part of bracket 4601 for receiving a bolt 4606. Yoke 4116 likewise has two spaced lugs 4608 formed integral therewith. Each lug 4608 has a hook portion 4612 formed at the end thereof, as seen best in FIG. 49. The hook portion 4612 of each lug 4608 engages with the bolt 4606. Each lug 4608 has a semi-circular recess 4614 provided in its surface for receiving a pin 4616.

The lower end 4617 of a lever 4618 is pivotally secured to the bolt 4606 which extends through the eye-lugs 4602. Lever 4618 is provided with a longitudinally extending recess 4620 for receiving a spring 4621. An elongated slot 4622 is also provided in lever 4618 through which the pin 4616 is inserted so that it can be supported by the lugs 4608. The spring 4621 is inserted in the recess 4620 through an aperture 4625 in end 4624 of the lever 4618. The aperture 4625 is closed by a screw 4626. The lower end of the spring 4621 is positioned on the pin 4616 by a pin 4630 which is secured to pin 4616 at a right angle thereto. The upper end of the spring 4621 rests against the screw 4626 in aperture 4625. Spring 4621 biases the lever 4618 to the position shown in FIGS. 48, 49 and 50. A mousetrap spring 4632, secured to the bolt 4606, is utilized to bias the lever 4618 in the stowed position, as indicated by the phantom lines in FIG. 49.

When the booster 500 is ignited and moves forward, the striker 4098 hits the upper end 4634 of the lever 4618, causing it to rotate about the bolt 4606. When this occurs, the yoke 4116 is pulled rearwardly by pin 4616 to shear the pin 4131 and release the sections 4103 and 4105 of the clamping ring 4036, as previously described.

As best seen in FIGS. 56, 57, and 58, there are six leaf springs 4125 provided approximately equidistantly spaced around the inner periphery of the clamping ring 4036. These springs 4125 insure that the clamping ring 4036 expands concentrically with respect to the missile aft ring 4096 and the thrust ring 4034 upon release of the sections 4103 and 4105 of the clamping ring 4036. These leaf springs 4125 are secured to the clamping ring 4036 by rivets 4128.

Six ring guards 4028, as shown in FIGS. 46, 47, 56, and 57, spaced equally around and attached to the forward end of the case 501 of the booster 500 by forwardly extending bracket members 4132 are employed to retain the clamping ring 4036 to the booster 500 after separation of the missile 100 therefrom. Members 4132 are secured to the booster case by bolts 4014. A ring guard 4028 is secured to each member 4132 by rivets 4136 as shown in FIG. 56.

As seen in FIGS. 46, 53, 54, and 55, two separation signal arms 4140 and 4142 are located 180.degree. apart on the horizontal centerline of booster flight. Each arm 4140 or 4142 is secured to a U-shaped bracket 4143 by a bolt 4144. Each bracket 4143, in turn, is attached to the booster case 501 by a bolt 4014. Each arm 4140 or 4142 protrudes from the forward end of the booster case 501 and mates with the missile 100 as will now be described.

Each signal arm 4140 and 4142 has a tongue portion 4146, which is received between the missile skin 286 and a spaced member 4148, which is secured to the missile skin by machine screws 4150. When the signal arms 4140 and 4142 are in position, the tongue portions 4146 thereof prevent the missile ignition system from being activated by preventing the biased elements 4152 in the microswitches 4154 from moving upwardly through apertures 4155 and 4156 in the missile skin 286 and member 4148, respectively, to complete electric connections. The signal arms 4140 and 4142 thus activate the missile ignition system as the booster 500 separates from the missile 100, as will be described. The microswitches 4154 are supported on the missile skin 286 by members 4160 and bolts 4162, and are contained in housing members 4164, with the latter being attached to the section 284 by welding or the like.

Referring now to FIGS. 46, 51, and 52, two indexing lugs 4170 and 4172 are used to position the booster 500 to the missile 100. These lugs 4170 and 4172 are located on the horizontal centerline of launch (viewed looking aft the booster 500) of the missile-to-booster thrust ring 4034 as best seen in FIG. 46. Two rectangular shaped notches 4174 (only one of which is shown in FIGS. 51 and 52) are provided in the missile aft ring 4096 for receiving the lugs 4170 and 4172. These lugs 4170 and 4172 are mounted in notches 4176 formed in the surface of the missile-to-booster thrust ring 4034. The lugs 4170 and 4172 are secured in the notches 4176 of the thrust ring 4034 by rivets 4178.

The two sets of forward and aft launching and handling shoes 4030 and 4032, and 4005 and 4007, as best seen in FIGS. 44, 45, 46, 47, 61, 62, and 64, are attached to the booster casing 501. The forward launching and handling shoes 4030 and 4032, respectively, are secured to the casing 501 with bolts 4180. The aft launching and handling shoes 4005 and 4007 are secured to the rear end 3999 of the casing 501 by bolts 4182 adjacent the forward end of the booster nozzle 4000 as best seen in FIGS. 59 and 62. The forward and aft launching shoes 4030 and 4005, respectively, ride in track assemblies 4412 and 4414 of the launcher 1000. These track assemblies 4412 and 4414, as will be described in more detail subsequently, support the missile-booster combination for the first 9 inches of its travel.

The propellant for the booster 500 is contained in a cylindrical grain 4184 pierced longitudinally by a plurality of spaced slots 4186, FIGS. 47, 59 and 60. A cast double base propellant is used in the booster 500. The casting powder is a modified nitrocellulose powder. The solvent used contains approximately 75 percent nitroglycerin, 24 percent plasticizer, and 1 percent stabilizer. The propellant is cast into a cellulose acetate cylinder (restrictor) which acts to prevent burning on the outer surfaces of the propellant grain 4184. Burning of the propellant is confined to the end of the grain 4184, and the surfaces of the longitudinal slots 4186 in the grain.

The propellant casting powder is comprised of nitrocellulose grains containing about 9 percent of a heavy metal compound, usually a lead compound. Lead stearate can be used. The addition of a heavy metallic compound modifies the burning of the propellant such that, in a given burning rate regime, the burning rate is relatively independent of ambient temperature. Burning of this type is termed "platonic" burning due to the "plateau" formed in the burning rate curve. The presence of the plateau is extremely desirable, as rocket thrust variation with propellant temperature is held to a minimum over the plateau range.

Whenever hot gases are forced to flow in a narrow channel at high velocity over a burning propellant surface, as they are in the booster 500, a phenomena termed "resonant burning" takes place. This type of burning results in extreme pressure fluctuations which may be severe enough to shatter the grain 4184 and rupture the case 501. In order to eliminate this phenomena, the resonance rod assembly 4026, FIGS. 47 and 59, is used. This assembly 4026 consists of resonance rods 4188 which are merely small diameter steel rods supported at the forward end by straps 4190 and 4191, and placed in the slots 4186 through which the gases flow. These rods 4188 effectively eliminate the resonance burning and result in a smooth burning pressure and smooth thrust output.

To prevent the booster case 501 from becoming overheated and weakened, the after end of the propellant grain 4184 is sealed to the booster case by resting against the shoulder 4006 machined on the booster nozzle 4000. A seam sealing compound or gasket 4004 is used between the shoulder 4006 and the case 501 to provide a gas tight seal. The grain 4184 is held down against the shoulder 4006 by springs 4192 of the grain immobilizer 4024. These springs 4192 are located between buttons 4193 located adjacent the inner surface of the forward head closure 4012 and pads 4194 placed adjacent to the upstream surface of the grain 4184. Telescoping rod members 4195 and 4196 are secured to the buttons 4193 and pads 4194, and the springs 4192 are placed about the rod members 4195 and 4196, as shown in FIG. 47, and the entire assembly is kept in operative position by straps 4197 positioned about annular rings 4341 and 4343. The springs 4192 provide a sealing force of between 775 and 925 pounds on the grain 4184 at -10.degree. F. These springs 4192 also prevent shifting of the grain 4184 during handling and transportation of the booster 500, and eliminate the need for seating the grain by elevating the launcher 1000 to maximum elevation before firing.

A typical curve of booster chamber pressure versus time is shown in FIG. 98. It will be noted that chamber pressure varies but little with changes in grain temperature. In FIG. 99, there is illustrated a typical curve of thrust in pounds versus time for the booster 500. The characteristics of the booster 500 for various temperatures are tabulated below:

40.degree. F. 77.degree. F. 125.degree. F. ______________________________________ Total Impulse 371,000 372,000 374,000 lb/sec lb/sec lb/sec Burning Time 4.10 sec 4.0 sec 3.70 sec (defined) Average Thrust 91,600 95,000 106,000 lbs lbs lbs Maximum Thrust 95,600 99,000 111,000 lbs lbs lbs Time at Maximum 0.2 sec 0.33 sec 0.25 sec Thrust Maximum Pressure 1085 1125 1290 lb/sq in lb/sq in lb/sq in Average Pressure 1030 1080 1220 lb/sq in lb/sq in lb/sq in ______________________________________

The igniter 4020 for the booster 500 is mounted on the forward head 4012, as seen in FIGS. 46 and 47. The electrical circuit for the igniter 4020 is illustrated in FIG. 100 (shown on the same sheet of drawings with FIGS. 56, 57, and 58). The igniter 4020 consists of a metal base 4302, which is a removable part of the forward head 4012, with an attached plastic cup 4304 attached thereto by means of an annular ring 4305. The ring 4305 has an inwardly extending flange 4307 provided integral therewith for mating with an outwardly extending flange 4309 on the cup. The annular ring 4305 is internally threaded and it is screwed on an externally threaded flange 4311 formed integral with the base 4302.

Cup 4304 contains an igniter charge 4306 and four electrical squibs 4308 connected together by leads 4310, 4312, 4313, 4314 and 4315, with leads 4312 and 4313 being connected to the metal base 4302, as shown in FIGS. 47 and 100. The center of lead 4310 is also connected to the terminal 4303 on the base 4302. A cone 4375 is utilized on base 4302 to insulate terminal 4303. A pin 4319 is utilized to pin the base 4302 to an annular flange 4321 formed integral with the forward head 4012.

A metal ring 4316 is provided for locking the igniter 4020 to the forward head 4012. This ring 4316 is scalloped in twelve equally spaced sections 4301 on its inner surface to provide therebetween circumferential sections 4317, and these sections 4317 fit over a corresponding number of circumferential lips 4318 formed on a flange 4323 which is formed integral with the igniter base 4302 to retain the igniter 4020 in the forward head 4012 when the booster 500 is armed. A retaining ring 4325, having a shoulder 4327 to cooperate with a flange 4329 on the ring 4316 maintains the metal locking ring 4316 in operative position.

In the safe position, the igniter lock ring 4316 is located so that the scalloped sections 4301 of ring 4316 are in alignment with the lips 4318 on the base 4302. A thin shear ring 4331 is positioned between the base scalloped sections 4317 and the lips 4318 to hold the base 4302 in position. If accidental ignition of the igniter 4020 should occur, the ring 4331 is sheared and the igniter base 4302 is ejected from its position in the forward head 4012, thus rendering the booster 500 non-propulsive. When the igniter 4020 is armed, the igniter lock ring 4316 is positioned so that the sections 4317 are located in alignment with the lips 4318 so as to retain the pressure load, thus maintaining the booster 500 in a propulsive state.

The igniter charge 4306 is comprised of a mixture of 300 grams of cannon black powder, 700 grams of black powder, and 500 grams of 0.012 inch thick trench mortar sheet powder which is cut into small segments. The four electrical squibs 4308 are wired in a series-parallel circuit and are, as previously indicated, internally grounded by leads 4312 and 4314 to the igniter base 4302. A lead 4362 is used as a jumper between the base 4302 and the booster forward head 4012.

The igniter 4020 is so mounted that, when it is in the safe position, the leads 4330 and 4332 for the squibs 4308 are disconnected from the firing lead 4540 and are shorted together to ground through the pole arm 4377 of a single pole double throw microswitch 4338.

To arm the booster 500, an arming handle (not shown) is inserted in the igniter arming device 4022 and rotated 90.degree.. As illustrated in FIGS. 46 and 47, this arming device 4022 has a gear sector 4344 mounted on its end which meshes with a second gear sector 4346 attached to the lock ring 4316 by screws 4350. This rotates the igniter 4020 22 1/2.degree., removes the short to ground on the squib leads 4330 and 4332 and connects the igniter 4020 to the firing circuit.

A fixed mechanical contact element 4352, as shown in FIG. 46, which is secured to the lock ring 4316 by a bracket 4356 and screws 4358, makes contact with a plunger 4360 on the switch 4338 to move pole arm 4377 to complete connection with contact 4379 and, at the same time, completes the firing circuit to lead 4540. The locking ring 4316, at the same time, locks the igniter 4020 in the forward head 4012. Switch 4338 is secured to the forward head 4012 by a bracket 4366 and screws 4368.

The launcher 1000, FIG. 61 comprises a central tubular column 4400 which is adapted to be fastened to the ground or a deck surface of a ship. A forward supporting leg 4402, and two aft supporting legs 4404 are provided which hold the column 4400 in a vertical position and also assist in restraining the column from lateral motion. A boom 4408, which is adjustable in elevation, is pivotally mounted on the upper end of the central column 4400. A hydraulic actuator 4406 is provided for elevating the boom 4408. Leveling means are provided on the legs 4402 and 4404 and are used to adjust the legs with respect to the central column 4400, thereby prestressing the launcher 1000 to withstand launching forces.

Mating forward and aft track assemblies 4412 and 4414 are carried by the boom 4408 and are arranged to cooperate with the forward and aft launching shoes 4030 and 4005, respectively, on the booster 500, as best seen in FIGS. 61 and 64, so that the missile 100 can easily be mounted on the boom. The mounting arrangement is such that the missile 100 is immediately freed of the launcher 1000 upon firing of the booster 500.

The launcher 1000 includes the central column 4400 which is supported by a base 4416 at its lower end. This base 4416 includes a plate 4418 which has an aperture (not shown) provided therein. This aperture is aligned with a second aperture in a block of concrete or similar material to which the base 4416 is attached by suitable means so that electrical cables, such as cable lead 4592, can be introduced into the interior of the column 4400 and boom 4408. A ball and socket arrangement (not shown) is mounted between the lower end of the column 4400 and the base 4416, respectively, to provide a pivotal connection between the base and column.

The cental column 4400 is held in vertical alignment and restrained from lateral motion by the arrangement of three supporting legs generally resembling that of a tripod and include the relatively long forwardly extending leg 4402 and two relatively short rearwardly extending legs 4404 (only one of which can be seen in FIG. 61). Because of the position of the missile 100 on the launcher 1000, the elements of the forward leg 4402 are made larger and structurally stronger than those of the rearwardly extending legs 4404. Each leg 4402 and 4404 is in the form of a simple truss. Legs 4404 each comprise spaced tubular members 4420 and 4422 interconnected by an end sleeve 4424, cross-member 4426 and gusset 4428. Leg 4402 comprises spaced tubular members 4430 and 4432 interconnected by a sleeve 4434, a cross-member 4436, and a gusset plate 4438, with all of these elements being structurally stronger than the corresponding members of legs 4404.

The tubular members 4430 and 4432, and 4420 and 4422, of the respective legs 4402 and 4404, are provided with flanges 4440 at their inner ends, which are removably attached by bolts 4442 to corresponding flanges 4444 carried on short tubular sections 4446, welded to column 4400, the weld joints being reinforced by gussets 4448.

The sleeves 4434 and 4424 at the outer ends of the legs 4402 and 4404 carry adjustable feet 4450. Each assembly includes a plate 4452 having a sleeve 4454 mounted thereon, the sleeve being reinforced by gusset plates 4456. A ball seat member (not shown) is mounted at the bottom of the sleeve 4454 for supporting a ball (not shown), which is fastened on the lower end of a post 4458 by a pin (not shown) and is retained in the sleeve 4454 by a retainer ring (not shown) screwed into the upper end of the sleeve 4454. This construction provides a ball and socket connection between each foot and post 4450 and 4458, respectively, at the outer ends of each of the legs 4402 and 4404.

The post 4458 has a threaded intermediate portion (not shown) which is screwed into a threaded aperture in a bushing (not shown), with the bushing being carried in the lower end of the supporting sleeve 4434 or 4424. A ring (not shown) is welded on the inside surface of the supporting sleeve 4434 or 4424 adjacent the upper end thereof to provide a shoulder, and an apertured guide disk, in which the upper portion of the post 4458 is journaled, is retained against the shoulder by an elongated cylindrical spacer extending between the bushing and disk. The upper end 4460 of the post 4458 is square to receive a wrench for adjusting the foot 4450 in the supporting sleeve 4434 or 4424.

Referring again to FIG. 61, the tapered boom 4408, which is of substantially rectangular cross section, is mounted on the upper end of the column 4400 by a pivotal connection generally indicated by reference numeral 4462. A plate 4464 having an aperture therein (not shown but used for the electrical cable 4592 brought up through the interior of the column 4400) is carried on the upper end of the column 4400, and it is attached to the column by a plurality of gusset plates 4466 which are welded to the column 4400 and plate 4464. The pivotal connection 4462 comprises a pair of spaced trunnion mounts 4468 attached to a plate 4470. This plate 4470, in turn, is secured by bolts 4472 to the underside of the boom 4408, nearer the larger end thereof. A pair of laterally spaced yokes or pillow blocks 4478 are bolted to the upper surface of the plate 4464. Pins 4480 are inserted in aligned apertures provided in the trunnion mounts 4468 and pillow block 4478 to complete the pivotal connection 4462.

The boom 4408 is comprised of an inner section 4482 and an outer section 4484 having flanges 4486 and 4488, respectively, which can be joined by pins, bolts or any desired means. The boom 4408 is fabricated from plate stock and it is hollow and substantially rectangular in cross section. Access apertures 4490 and aperture 4492 are provided in the boom 4408, the former adjacent the inner end of the boom and the latter adjacent the outer end thereof.

Forward and aft track assemblies 4412 and 4414, which cooperate with the forward and aft booster shoes 4030 and 4005, respectively, as best seen in FIGS. 61, 62 and 64, are supported on section 4484 of boom 4408, as will now be described. The track assemblies 4412 and 4414 are removably mounted on a plate 4496 which is attached by bolts 4498 to the underside 4494 of boom section 4484, with suitable spacers 4499 being inserted between the plate 4496 and the boom underside 4494.

The plates 4496, FIGS. 62 and 63, are provided with a notch 4502 along their longitudinally extending edges which notches are adapted to receive eye-bolts 4504. These eyebolts 4504 are pivotally mounted on bolts 4506 journaled in spaced yokes or lugs 4508 projecting from opposite sides of a box structure 4510. This box structure 4510 consists of a channel 4511, which has a plate 4513 welded to the top thereof, as best seen in FIG. 64. The aft end is similarly provided with spaced yokes or lugs 4508 for bolts 4506 which pivotally mount eye-bolts 4504. The foward track assembly 4412 includes two forward tracks 4520 which are welded to the box structure 4510, with each track being provided with a groove 4516 into which a lip 4518 in the corresponding forward booster shoe 4030 projects.

The aft track assembly 4414 is generally similar to the forward track assembly 4412 except for the details of the aft track 4514. In the aft track assembly 4414, the single launching shoe 4005 is T-shaped in cross-section and is received, as shown in FIGS. 62 and 64, in a correspondingly shaped recess 4526 in the aft track 4514. This aft track 4514 is welded to the underside of channel 4511. Tapered pins, such as pin 4530, shown in FIG. 62, are utilized to position the box like structure 4510 with respect to plate 4496. Stop 4532, as shown in FIG. 63, is provided in the recess 4526 near its aft end to prevent rearward movment of the booster 500 when it is mounted on the launcher 1000.

In FIGS. 46, 64 and 65, the details of the electrical connections to the igniter lead 4540 are illustrated. As shown in FIG. 65, a flat groove 4542 is provided in the upper surface of left hand forward launching shoe 4030. A strip 4544 of insulating material is placed in the groove 4542, and a strap 4546 of conducting material is positioned on top of the strip 4544.

The ends of the strap 4546, as shown in FIGS. 46, 62 and 64 are shaped to fit over the front and rear surfaces of shoe 4030 and are attached thereto by screws 4543, only one of which is shown. As indicated, the strip 4544 of insulating material is similar in shape to strap 4546, and it separates the strap 4546 from the shoe 4030. A shouldered bushing (not shown) of insulating material is provided between the strap 4546 and the screws 4543. A terminal lug 4545, insulated from the shoe 4030 and screw 4543 by the shouldered bushing, is held in electrical contact with strap 4546. This terminal lug 4545 is carried on one end of the igniter lead 4540 from the igniter 4020.

As shown in FIGS. 64 and 65, an insulating block 4548 of phenolic material is secured to the side 4550 of the box structure 4510 by suitable means, such as screws 4552, which pass through apertures 4549 in the block 4548 and which are received in threaded apertures 4554. Block 4548 is provided with a vertically extending aperture 4556 which is of reduced diameter for the lower half of the block 4548 and of enlarged diameter for the upper half thereof.

A metal rod 4558 is inserted in the aperture 4556 in the block 4548, and then a coil spring 4560 is inserted about the upper end thereof in the enlarged diameter portion of the aperture. An insulating sleeve 4562 is then inserted about the upper end of the rod 4558, and it is pinned to the rod 4558 by a pin 4564. Coil spring 4560 thus biases the rod 4558 upwardly by pushing against the sleeve 4562 and the shoulder 4569 in the aperture 4556. This rod 4558 provides an electrical contact 4566 on its upper end. An enlongated slot 4568 is provided in the rod 4558 near its lower end, and a pin 4570 is inserted through an aperture 4571 in the block 4548 and through slot 4568. The pin 4570 allows limited vertical movement of the rod 4558 and retains the rod 4558 in the block 4548 when the launcher is unoccupied.

Electrical connection between the rod 4558 and the metal strap 4546 is provided by a strap 4574 which has one end 4576 thereof inserted into an aperture provided in the lower end of the rod 4558, and whose other end 4578 has a contact element 4580 mounted thereon. This element 4580 rests on the metal strap 4546.

Plate 4496 is provided with an aperture 4581 into which there is inserted a sleeve 4582 formed of insulating material. A metal rod 4584 is inserted in the sleeve 4582, and it is provided with an enlarged end contact 4586 for mating with contact 4566 on the upper end of rod 4558. The upper end of rod 4584 is threaded to receive two nuts 4588. A terminal lug 4590 from cable lead 4592 is positioned between these two nuts 4588. Cable lead 4592 is passed through an aperture 4594 in the side wall 4596 of boom 4408. This cable lead 4592 is passed through the boom 4408 and column 4400. It will be discussed more subsequently in connection with the igniter firing circuit for the booster 500.

Terminal lug 4590, cable lead 4592, and nuts 4588 are contained in a housing 4591, which is secured to the side wall 4596 of the boom 4408 by bolts 4593.

Referring to FIGS. 63 and 64, the forward track assembly 4412 is provided with a removable safety stop 4640 to prevent displacement of the missile 100 from the boom 4408. The removable safety stop 4640 is formed of heavy rod and is received in a suitable aperture 4641 provided in the forward track 4520 and it can be readily withdrawn from the path of the booster shoe 4030 by means of a circular handle 4642.

The release lever striker 4098 is mounted on the right side of the launcher boom 4408, looking in the direction of the missile 100. An outwardly extending channel 4644 is secured to the box structure 4510, as best seen in FIGS. 63 and 64, by suitable means such as bolts 4651. A second channel 4646 is, in turn, secured to the outwardly extending channel 4644.

This second channel 4646 is located substantially parallel to the box structure 4510 but is spaced therefrom. A gusset plate 4648 is used is strengthen the channels 4644 and 4646. A pair of spaced lugs 4650 are welded to the forward end 4654 of channel 4646. The release lever striker 4098, in the form of a bolt, is supported by these lugs 4650, and it is positioned so as to strike the upper end 4634 of lever 4618 when the booster 500 is ignited and moves forward, as previously described, thereby releasing the sections 4103 and 4105 of the clamping ring 4036.

As best seen in FIG. 61, the hydraulic actuator 4406 is used to elevate the launcher boom 4408 to the desired firing angle. The actuator 4406 is of a conventional type and it includes a cylinder 4660 which is pivotally mounted on the column 4400 near its lower end. This pivotal mounting includes a plate 4662 attached to the column 4400 and having apertured lugs 4664 integral therewith. A hinge 4666 is carried on the lower end of the cylinder 4660 and a pin 4668 is inserted through aligned apertures provided in the lugs 4664 and the hinge 4666. The piston rod 4670 of the hydraulic actuator 4406 is pivotally attached at its upper end to the inner end of the boom 4408. A pair of spaced stop rods (not shown) located on either side of, and parallel to the piston rod 4670, are likewise pivotally attached at their upper ends to the boom.

The pivotal connection between the piston rod 4670 (as well as the stop rods) and the boom 4408 includes a plate 4674 attached by bolts 4676 to the end of boom section 4482. The plate 4674 has three pairs of suitably spaced apertured lugs 4678 formed integral therewith. The upper ends of the piston rod 4670, as well as the two stop rods, are provided with sleeve bearings 4680, and a shaft 4682 is inserted through aligned apertures in the bearings and the lugs, with the shaft being retained in position by snap rings (not shown) provided adjacent its outer ends.

The stop rods (not shown) are slidably carried in spaced apertures provided in a stop block 4648 on cylinder 4660. A fully comprehensive description of the elevating mechanism and attendant details is provided in the patent application of Arthur H. Miller, Jr., Ser. No. 437016 for a LAUNCHER, filed June 15, 1954.

The operating mechanism for the hydraulic actuator 4406 includes a pump driven by an electric motor, a valve, and a tank for the hydraulic fluid, and is mounted in a housing 4686 on the lower member 4432 of the forward leg 4402. A control handle (not shown) is provided on the housing 4686. Since these elements are conventional, no detailed description is given herein.

A pair of hydraulic leads (not shown) interconnect the pump housing 4686 and the hydraulic cylinder 4660, with one lead being connected to the lower end of the cylinder and the second lead being connected to the upper end thereof. Flexible sections (not shown) are provided in the hydraulic leads to permit bending as the cylinder 4660 moves about its lower end. A flow limiter is inserted in the upper hydraulic lead to limit the flow out of the cylinder 4660 in the event a power failure stops the pump. The flow limiter prevents a loaded boom 4408 from moving rapidly to its lowermost position.

An indicator (not shown) is mounted on the column 4400 for assisting in vertically aligning the column.

As shown in FIG. 61, a panel 4690 is also provided on the central column 4400 for mounting such controls as may be needed, and an oil filler tank 4692 is provided on the column 4400 for filling the actuator cylinder 4660.

It is belived that the general mode of operation of the launcher 1000 will be apparent from the foregoing detailed description. To facilitate mounting the missile 100 and its booster 500 on the boom 4408, the forward and aft track assemblies 4412 and 4414 and box structure 4510 are removed and are mounted on the booster 500, the forward booster shoes 4030 being received in the grooves 4516 of the forward track assembly 4412 and the aft booster shoe 4005 being received in the T-shaped recess 4526 in the aft track assembly 4414. The missile 100 and the booster 500 are then raised to a position near the boom 4408 and in alignment therewith, so that the eye-bolts 4504 can be swung into their respective notches 4502.

With the missile 100 in position, the central column 4400 can be aligned vertically by adjusting the feet 4450, if necessary, and prestressing of the launching structure can also be accomplished, if desired, by adjusting the feet 4450 with the central column 4400 fixed. The stop rings are then adjusted along the stop rods to positions corresponding to the desired angle of elevation of the boom 4408, and the hydraulic actuator 4406 is energized to raise or lower the boom 4408 to the required angle. Firing is accomplished by the ignition of the booster 500 from a remote point, as will be described subsequently, through the medium of an electric current applied through the leads 4592 and 4540. Because the forward and aft tracks are extremely short, the missile 100 and booster 500 are quickly freed from the launcher 1000.

The missile of the present invention may utilize a conventional warhead, or a warhead of the type described in U.S. Patent application, Ser. No. 590,077, filed June 7, 1956, E. L. Nooker and M. L Kempton, Inventors (application assigned to U.S. Government). Such a warhead is shown in FIGS. 66-68 of the drawings.

The fuze 182 for the missile 100 is a fixed angle, narrow beam, microwave, influence type proximity fuze, having a range of approximately 140 feet. As shown in FIGS. 1, 4, and 43A, the fuze 182 is located in the forward annular nose section 104 of the missile 100 below the fuze compartment cover 190. Six fuze antennae 184 are located in section 104 to provide the radiation pattern at a fixed angle with the longitudinal axis of the missile.

The fuze 182 is designed so that its radiation pattern coincides with the dynamic burst pattern of the warhead 194. The fuze 182 transmits microwave energy, which, when the missile 100 is sufficiently close to the target 2000, is reflected and received by the fuze to produce a voltage which actuates the detonators 4800, which, in turn, detonate the warhead 194.

In order to insure operation of the fuze 182 on the correct target 2000 and to insure maximum damage from the blast pattern of the warhead 194, steering intelligence information is fed to the fuze, thus enabling it to trigger at the optimum time in flight for any approach to the target 2000.

The safety and arming device 132, as illustrated in FIGS. 1 and 101, renders the warhead 194 and/or self-destruct charges inoperative until such time as it is deemed safe to detonate these charges. The arming of the self-destruction charge will not be covered since the operation is identical to arming of the warhead 194. The safety and arming device 132 will not allow the warhead 194 to arm until an electric circuit is provided for the detonate signal from the fuze 182, and a flame path is provided between the detonators 4800 and the booster for the warhead 194.

The operation of the safety and arming device 132, in order of events, will now be explained in conjunction with FIG. 101. Two selector switches 4900 and 4902 on the safety and arming device 132 are adjusted and locked before launching of the missile 100. Switch 4900 is utilized to select either the warhead 194 or exercise head 196; while the other switch 4902 is used to select the arming time, either 9 or 18 seconds after launching of the missile 100.

The next event in the operation of the safety and arming device 132 is to provide an intent to launch signal to the safety and arming device. This signal is received when internal power is applied to the missile 100 approximately 45 seconds prior to launching of the missile 100. This applies -330 volts through a resistor 4904 to contact 4906 on the safety and arming device connector 4904, and also provides -330 volts on the contacts 4910 and 4912 of a G operated switch 4914.

At zero time, the missile 100 is launched. When the missile 100 attains a 10g acceleration, the switch 4914 moves to position 4916. This applies -330 volts to a resistor 4918 through the contacts 4910 and 4920, thus providing a charge path for a capacitor 4922. When the voltage on capacitor 4922 rises to a high enough potential, gas discharge tubes 4924 and 4926 fire, thus causing a current flow through the squibs 4928 and 4930 on the explosive dimple motor 4932. The RC time constant of resistor 4918 and the capacitor 4922 is selected so that it will require 2.25 seconds for the gas discharge tubes 4924 and 4926 to fire. The -330 volts is also applied to a capacitor 4934 charging it very quickly to -330 volts. It should be noted that unless the 10g acceleration of the missile 100 is sustained for 2.25 seconds, the capacitor 4922 will not charge to a potential sufficient to fire the tubes 4924 and 4926.

When the current flows through the squibs 4928 and 4930 on the explosive dimple motor 4932, the squibs 4928 and 4930 fire and the motor drives the switch 4914 to position 4936 and locks it in this position. When the switch 4914 locks in position 4936, it provides a discharge path for the capacitor 4934, through contacts 4938 and 4940 on switch 4914 through resistors 4942, 4944, and a capacitor 4946. Consequently, capacitor 4946 starts charging. When the potential across capacitor 4946 rises high enough, gas discharge tubes 4948 and 4950 fire and cause current to flow through squibs 4952 and 4954 on an explosive bellows motor 4956.

The RC time constant of capacitor 4934, resistors 4942 and 4944, and capacitor 4946, is selected so that the time for the gas discharge tubes 4948 and 4950 to fire will be 15.75 or 6.75 seconds after the G switch 4914 has been in position 4936. As stated previously, this time was determined by the setting of selector switch 4902 before launching of the missile 100.

The motor 4956 is mechanically connected to a plate 4958, and switches 4960, 4962, and 4964. When the squibs 4952 and 4954 of motor 4956 fire, the motor drives the plate 4958 to the position that connects the detonators 4800 electrically to ground at point 4966 and to the detonate signal lead at point 4968.

Also, the detonate signal path is completed from the fuze 182 by the rotation of the switch 4962 and plate 4958, when rotated, completes the flame path from the detonators 4800 to the booster of the warhead 194. This completes warhead arming. It should be noted that switch 4900 determines whether the detonate signal will be sent to the warhead 194 or to the exercise head 196.

An exercise head assembly 196 is used in lieu of the warhead 194, as shown in FIG. 43C, to obtain flight data for research and development. This also permits operation against an actual target 2000, such as an aircraft, without destroying it.

The exercise head assembly is composed of two sections 4840 and 4842. Section 4840 contains the telemetering equipment 206, while the other section 4842 contains various telemetering components, several (two) spotting charges, and a self-destruct charge.

The two spotting charges, which are used to simulate warhead detonation, are exploded by igniters. Actuation of the igniters is controlled by the fuze 182. Each spotting charge produces a smoke puff with an intensity of 8 megalumens. The self-destruct charge is fired upon command from the steering intelligence system 212. Both the spotting charges and the self-destruct charge are armed by the safety and arming device 132.

The missile weapon system is comprised of the equipment necessary to effectively deliver fire upon high speed hostile aircarft with the missile 100, as previously described, at ranges up to 100,000 yards and at altitudes up to 60,000 feet. The missile weapon system includes search and height finding radars, identification equipment, target designation systems, the missile fire control system, the handling and launching equipment, the missile 100 with its various systems as previously described, and the personnel and supply systems necessary to insure proper utilization of the missile 100.

The missile weapon system can be used on land or aboard a missile ship, and when used aboard the latter, it functions to provide long range, high fire-power anti-aircraft defense to a task group, supplementing conventional weapons. The missile ship, of the light cruiser type, mounts two dual launchers 5400, only one of which is shown in FIG. 103. One of the launchers 5400 is located in the forward part of the ship while the other is mounted in the aft portion thereof and on the main deck. A typical missile ship will carry missiles, stowed in magazines 5401, as shown in FIGS. 104 and 105.

The missiles 100 are received aboard the missile ship, the hydraulic system 220 pressurized, and fuel tanks 138 filled. Missiles 100 and boosters 500 are shipped in airtight cans. Immediately on arrival aboard the ship, the missiles are decanned, mated, and taken below deck where they are horizontally stowed. Wings 112, 113, 114, and 115, and fins 116, 117, 118, and 119 are not attached until just prior to firing of the missile-booster combination.

In the vertical stowage and loading plans, mated, fueled and charged missile-booster combinations, with the warhead 194 and fuze 182 installed, are placed on stowage arms in racks capable of holding a plurality of missile-booster combinations each. At least four racks are required to serve one dual launcher 5400. The launcher 5400 is located between them, and above the tops of the stowage arms. Four blast doors are located in the deck immediately beneath the launcher 5400. These doors allow passage of the missile-booster combination and the rammer loader. It can be readily seen that the compartmentation and watertight integrity problems in this type of installation become immense. It requires a space beneath each dual launcher 5400 of approximately 28 feet wide, 34 feet long, and 35 feet high. This height will extend from the first platform deck up to the first (01) deck and second (02) deck, respectively.

Mounted on a cylindrical platform are nine stowage arms. These arms are supported at the top and bottom, and are arranged to travel independently around the cylindrical platform. In this position, only one set of fins and wings for each missile 100 can be installed permanently. Space will not permit the installation of the other pair of fins and wings on the missile, until the missile-booster combination reaches the final position before being rammed up to the launcher 5400. Also, a rather large strike down area is required.

Warmup power can be applied to all of the missile-booster combinations on the stowage arms. A Go-No-Go tester is used to check out the missile 1100. However, here if a missile-booster combination fails this test, it can be easily cycled around to the rear of the cylinder and a new missile-booster combination brought up for test.

In horizontal stowage and loading, on the other hand, mated, fueled and charged missile-booster combinations are stored below the ship deck in horizontal racks. Each of the stowage racks is a watertight compartment. The missile-booster combination is placed on a light dolly which, in turn, is mounted on rails. The rails run athwartships and terminate at the hoist. The hoist is provided with short rails on it which will mate with the rails from the racks. The missile-booster combination will be rolled out to the hoist and onto the hoist to the standby position.

When a missile-booster combination is needed in the wing and fin assembly room, the hoist is run up to this room and the missile-booster combination is transferred to the ram assembly. Here the wings and fins are installed after the final Go-No-Go test has been passed. In the floor of this room, there will be located inside blast proof doors. It is through these doors that the missile 100 will be passed from the hoist to the ram assembly. At the front of this room, and directly behind the launcher 5400 it will serve, another pair of blast proof doors is provided. These two pairs of doors are so interconnected by interlocks that only one set can be open at any one time. This is to prevent any accident in one room being spread into any other room.

A retractable rail, operating on an overhead pivot, is located directly behind the blast proof doors. This rail is swung out when the outside blast proof doors are open to make the rail connection from the wing and fin assembly room to the launcher 5400. When this rail is in position, the rammer assembly is cycled to move the missile-booster combination onto the launcher arm. Because of the close tolerances in the shoes that hold the missile-booster combination to the launcher arm and the ram assembly, the launcher 5400 is locked in the loading position and not simply held there by action of the elevate and train servos.

After the missile-booster combination has been transferred from the dolly on the hoist to the ram assembly, the dolly can be returned to the strike down area to await the arrival on board of the resupply missile-booster combination.

If the warhead 194 has not been previously installed during shipment, a separate warhead hoist is utilized to bring the warheads 194 up from their magazine to the wing and fin assembly area for installation in the missile-booster combination. The warhead 194 is in two halves, each semi-circular in shape as previously described, and these halves are bolted around the missile 100 prior to firing. The fuzes are installed at this same time.

The fuel and hydraulic systems 230 and 220, respectively, are designed to be tight in order that they can be charged at the time the missile-booster combination is originally mated, and they will remain in this condition at all times during storage. The missile batteries are also charged and installed at this same time.

The main advantage of this type of horizontal storage and handling system is that the ship as a whole can be made much more watertight, and the system of compartmentation can be better carried out in the missile handling compartments of the ship. The actual cubic space requirements are approximately the same for both the horizontal and vertical systems for storage of the missile-booster combinations. However, some extra equipment will be required, such as the dollies and the hoist.

The launchers 5400 are of the dual type, capable of firing salvo fire, with the two missiles 100 being separated in space by about 2 seconds. These launchers 5400 are located in the normal main battery positions, that is, two forward and two aft, on the ship centerline. The No. 1 and 4 launchers will be on the main deck, while the No. 2 and 3 will be on the first (01) deck of the ship.

The launchers 5400 are trainable through 360.degree. and can be elevated up to 90.degree.. Cut-out circuits will be employed to prevent the firing of missiles 100 into the superstructure of the firing ship.

To load a missile-booster combination onto the arms of the launcher 5400, it is only necessary to train the launcher 5400 around to either 0.degree. or 180.degree. relative, and either elevate or depress to 90.degree. or 0.degree., open the blast doors, and then operate the rammer. In the vertical load system, the launcher 5400 must be at 90.degree. of elevation to load it, while in the horizontal load system it must be 0.degree..

In the vertical load system, any launcher 5400 can be loaded when at either 0.degree. or 180.degree. relative, while in the horizontal load system previously described, the two forward launchers 5400 must be at 0.degree., and the aft launchers 5400 must be at 180.degree. relative. This may tend to slow down the rate of fire for the horizontal load system.

In both stowage systems, the missile-booster combination is carried on a loading arm, which is attached to the rammer. This loading arm is identical to the launcher arm, and when it is in physical contact with the end of the launching arm, the rammer pushes the missile-booster combination off of the loading arm onto the launching arm. A one way catch snaps shut after the missile 100 passes over it on the launcher arm. This will prevent the missile-booster combination from falling off of the launcher 5400 during the retraction of the rammer and during the elevation and training of the launcher.

As shown in FIG. 102, the ship will have two independent installations for firing missile salvos. Each installation 5402 and 5404, called a defense unit, consists basically of one launcher 5400, one launcher computer 5406, two tracking and illuminating radars 5408 and 5410, two predictors 5411 and 5413, two beam programmers 5415 and 5417, two guidance transmitters 5412 and 5414, a control console 5434, and a data receiver and transmitter 5432.

A missile direction center 5418 is utilized to coordinate the actions of the defense units 5402 and 5404. This direction center 5418 keeps track of all targets 2000, and designates them to the two defense units 5402 and 5404.

If it is assumed that a target 2000 has been detected by a combat information center 5420, it is evaluated and assigned to the missile direction center 5418 as a target. The target 2000 is designated to a defense unit, say defense unit 5402, which issues a warmup order to the launching system and commences tracking the target 2000.

The launching system automatically performs the sequence of operations described below. It starts the warmup of the missile 100; and it transfers missiles 100 from the magazine 5401 to loader area 5422 where wings 112, 113, 114, and 115, and fins 116, 117, 118, and 119, are attached to missile 100, and the fins 256, 257, 258, and 259 are attached to the booster 500. In addition, it rams the missile salvo onto the launcher 5400; and then it elevates and trains the launcher 5400 to the proper position as determined by the launching computer 5406.

The launching computer 5406 also determines the initial train and elevation of the assigned guidance transmitter 5412 or 5414 for missile capture, correcting for the effects of gravity drop, parallax, and wind. As soon as the guidance transmitter 5412 or 5414 is synchronized to the proper position, the firing sequence is initiated. At the instant the firing command is given, launcher orders are frozen to the true space values then existing, and held until the second missile 100 of the salvo is fired. The beam position of the guidance transmitter 5412 is also held fixed in true space for the first 6 seconds after firing to facilitate beam capture by both missiles 100. As soon as the second missile 100 is clear of the launcher 5400, the launcher slews to the reload position.

How the missile weapon system fits into the overall air defense scheme will now be explained. All raids approaching a task group will be evaluated on the task group command level and certain targets will be assigned to the ship. Aboard the ship these target assignments are received in the missile direction center 5418, which, in turn, designates the targets 2000 to the two defense units 5402 or 5404. Each defense unit 5402 or 5404, operating in the dual-simplex mode of operation, is capable of simultaneously controlling two missile salvos. The function of each part of the weapon system, illustrated in FIG. 102, is illustrated below.

The combat information center 5420 detects and identifies all targets 2000; and it continuously tracks all targets 2000. In addition, this center 5420 correlates all available data about each target 2000 which will assist the evaluation and weapon assignment operations. It also evaluates the relative threat of the various targets 2000; selects the targets 2000; and there it distributes the targets 2000 among the missile ships.

The missile direction center 5418 designates assigned targets 2000 to the two defense units 5402 and 5404. Targets 2000 are assigned successively in priority sequence until each defense unit 5402 or 5404 is loaded to capacity. Subsequent targets 2000 are assigned in priority sequence as the tracking channels are released from their prior targets. In designating targets to a defense unit 5402 or 5404, consideration is given to the tracking radar and guidance transmitter blind zones.

The missile defense unit 5402 or 5404 is the combination of equipment required for firing the missile 100. Each defense unit 5402 or 5404 operates on its assigned targets 2000 independently of the other defense unit, and it is capable of autonomous operation in the event of battle damage. The defense unit controller directs and coordinates the operation of the guidance transmitter 5412 or 5414, tracking radars 5408 and 5410, and the launching system.

In response to the orders received from the missile direction center 5418, the controller signals the launching system as to the missile 100 which is to be fired. The controller then selects the tracking radar 5408 or 5410 and guidance transmitter 5412 or 5414 which is to be used against a particular target 2000 after considering blind zone regions with respect to the present and predicted intercept positions of each target.

Firing is normally automatic after receipt of the designation data from the missile direction center 5418 as soon as all the conditions required for successful intercept are satisfied. Among these are targets 2000 within range; missiles 100 warmed up on the launchers 5400; the launcher 5400 pointed correctly; and the guidance transmitter 5412 or 5414 positioned for capture. The defense unit commander will have the firing circuit under his control so that he can prevent automatic firing if such is called for by his overall knowledge of defense unit status and safety conditions. In addition he will be able to bypass interlocks for emergency firing.

Each defense unit 5402 or 5404 contains two tracking and illuminating radars 5408 and 5410 which provide target position and rate data to the predictors 5411 and 5413 and beam programmers 5415 and 5417, and provide target illumination for the homing guidance sub-system of the missile 100. The radars 5408 and 5410 consist basically of two separate radar systems, one for acquisition and one for tracking. Both of these operate simultaneously and transmit through the same antenna lens.

The acquisition system supplies a narrow beam that scans a volume in space approximately 5.degree. wide, 5.degree. high, and 10,000 yards deep in range. The acquisition bean assists in the process of locking the track beam on a designated target. The track system, which is of the simultaneous lobing type, supplies a narrow 1.6.degree. beam for automatic tracking. Using both systems, it is possible to survey the space surrounding a target 2000 while the target is being tracked.

The two predictors 5411 or 5413 control the automatic sequence of defense unit operations. Each predictor 5411 or 5413 is permanently associated with a particular channel of the target tracking system. Each predictor 5411 or 5413 continuously computes the three coordinates of the predicted target intercept point; the missile time of flight to predicted intercept point; and the time duration required for the beam program.

The operations listed below are automatically initiated by the predictor 5411 or 5413 on the basis of these computations. The loading of the launcher 5400 is initiated about 50 seconds before predicted firing; and then the firing of the missile 100 is initiated. The tracking radar 5408 or 5410 is then directed to acquire its next target 2000 when the predicted intercept time for the present target is reached. The guidance transmitter 5412 or 5414 is released for use with the next salvo fired.

The launching computer 5406 performs three functions in controlling the relationship between the target 2000, launcher 5400, guidance beam and missiles 100. The launching computer 5406, using predictor data, generates the train and elevation orders for setting the launcher 5400 to its firing position. For all targets 2000 having a predicted intercept point beyond 50,000 yards, the launching elevation angle with respect to the horizon will be approximately 40.degree.. For closer intercept ranges, the required setting varies in 5.degree. steps between 25.degree. and 50.degree.. At the instant the firing command is given, the launching computer 5406 freezes the launcher position to the true space values then existing and holds it fixed in this position until the second missile 100 of the salvo is fired.

The launching computer 5406 also determines the initial train and elevation angle of the guidance transmitter beam. Since the missile 100 is not guided during the boost phase, but is only attitude stabilized, it is necessary that the guidance transmitter beam be positioned properly to capture the missile 100 when its guidance control system 214 is activated. The initial train and elevation positions required of the guidance transmitter beam are obtained by adding the effects of gravity drop, parallax and wind to the angles of the launcher 5400. The guidance transmitter beam position is held fixed in true space for the first 6 seconds after firing, and the beam position programming is initiated at this time.

The launching computer 5406 also will measure the cross traverse displacement of the launcher 5400 at the instant the missile gyroscopes are uncaged and will generate orders for corresponding displacements of the polarization and modulation axes of the guidance transmitter assigned to that salvo. These guidance transmitter axes will be continuously modified during missile flight to compensate for changes in cross traverse position resulting from both beam programming and ship motion.

The beam programmer 5415 or 5417 programs the position of the guidance radar beam so that the missile 100 follows an up-and-over trajectory. The program used for the elevation position of the guidance transmitter beam is: ##EQU3## where E.sub.G is the present elevation angle of guidance transmitter 5412 or 5414;

E.sub.t is the present elevation angle of target 2000 (from tracking radar 5408 or 5410);

E.sub.o is the elevation angle of the guidance transmitter 5412 or 5414 at the instant of firing the missile salvo;

T.sub.gp is the time from start of beam program to present time (F (t) potentiometer); and

T.sub.gi is the time from start of program to predicted intercept time (from predictor 5411 or 5413).

It can be seen that when T.sub.gp = 0, E.sub.G = E.sub.o ; and when T.sub.gp = 1.2 T.sub.gi, E.sub.g = E.sub.t, showing that the guidance beam is programmed into coincidence with the tracking radar beam at the end of the program. A similar program is used in the azimuth.

Each defense unit 5402 or 5404 includes the previously mentioned guidance transmitters 5412 and 5414 for the concurrent control of two missile salvos. These units produce the beams along which the missiles ride. The beam is conically scanned.

The launching system operates under the overall control of the predictors 5411 or 5413 of the defense unit controller. Thus, missile type selection, initiation of launcher loading, and initiation of firing are all under the direction of the defense unit controller.

Except for the launcher 5400, the entire launching system is enclosed in an armored deckhouse. As previously mentioned, automatically operating doors separate the launcher 5400 from the loader area and another set of doors separates that area from the magazine. It takes approximately 50 seconds for a missile salvo to be hoisted to the loader area, have aerodynamic surfaces installed, be rammed onto the launcher 5400, and be elevated to firing position. Immediately after firing, the launcher 5400 returns to the horizontal reload position.

The general operation of the overall missile system will now be described. In response to signals generated in the radar 3408 (or 5410) and launcher controller 5434 in the missile direction center 5418, the launcher 5400 (with two missile-booster combinations on the launcher arms) trains and elevates to the firing bearing and quadrant elevation. The bearing is very nearly that of the target 2000, while the elevation varies with the range, being high for long range and low for short range.

The same computer 5406 through the beam controller 5415 trains the midcourse guidance transmitter 5412 so that its axis is in coincidence with the predicted missile position at the end of boost. This midcourse guidance transmitter 5412 does not track the missile 100, but it transmits pulses of RF energy in the beam 2024. This beam 2024 is programmed in elevation to give the Up and Over trajectory that is desired for long range and fuel economy, coincident with target tracking.

When the firing circuit is closed through the pair of microswitches 4154, shown in FIG. 53, a programmed series of events takes place. The missile 100 is switched to internal power, a backscratcher (not shown) (a missile launcher electrical receptacle is shown on line 1, FIG. 43D through which the external power is applied) for furnishing external power is lifted and the booster 500 is fired. This booster 500 will deliver 370,000 lb/sec of impulse, which is equivalent to about 93,000 pounds of thrust for about 4 seconds. This thrust is sufficient to bring the missile 100 to near its designed velocity, 2000 ft/sec, at an altitude of between two or three thousand feet, depending upon the firing quadrant elevation.

When the thrust from the booster 500 has been expended, its greater aerodynamic drag causes it to slow down faster than does the missile 100 and it falls away from the missile. This action completes the separation circuits in the missile 100. This activates the fuel control system 230 and the beam riding guidance sub-system 2012.

During the boost phase of aerial flight of the missile-booster combination, excessive dispersion of the point of separation caused by wind, and aerodynamic or jet misalignments can be expected. At the present time, this dispersion is reduced by the use of the large booster fins 256, 257, 258, and 259.

At separation of the missile 100 and booster 500, the flares in the combustor 108 are ignited and the ramjet engine takes over the propulsion of the missile. This engine burns the fuel in the annular can combustor 108 to give thrust and keep the missile 100 at its designed velocity. As previously mentioned, two burner sections are present in the combustor namely, a main burner and a pilot burner. The purpose of the pilot burner is to stabilize the burning in the main combustor 376 and to ignite the main burner if the main fuel flow shuts off due to missile velocities exceeding the design velocity. The main fuel regulator 380 regulates fuel flow to the combustor 376 to hold the missile velocity at 2000 ft/sec.

The tank 442 of nitrogen aboard the missile, operating through a pressure regulator unit 444, pressurizes the neoprene bladders 438 inside the fuel tank 138 to a pressure equal to ram air pressure plus 28 pounds per square inch. The nitrogen system provides a positive head of fuel to the pump 440 to overcome acceleration forces at launch, prevent cavitation and fuel boiling at high altitudes.

Ram air picked up by the scoop 772 inside the air duct 266 goes through the air turbine 770 which turns the fuel pump 440. This assures a constant flow of fuel to the fuel regulator valves which control its flow to the pilot and main combustors.

The hydraulic system 220 is also pressurized by means of pump 1390 which is operated by the air turbine 1282. This turbine 1282 also runs off of ram air, in a similar manner to the fuel pump 440. The hydraulic oil flows from the previously pressurized accumulators 122 and 1402, to the actuator valves that move the wings 112, 113, 114, and 115, then back to a sump tank 1396 where it waits to be recirculated by the pump 1390.

It is assumed that at the end of the boost phase of aerial flight, the missile 100 will be in such a position in the beam 2024 of the guidance transmitter 5412 that the beam rider receiver 2042 can determine the location of the missile in this beam and commence to steer the missile along it. However, if for some reason the orientation of the missile 100 were such that the missile were upside down and high in the beam, the receiver 2042 would determine that the missile was high in the beam and send a down-signal to the wings 112, 113, 114, and 115. But since it is upside down, a down-signal would make the missile 100 go up and cause it to get farther out of the beam 2024. For this reason, the output of the beam rider receiver 2042 is shorted to ground for 1.1 seconds. This is sufficient time for the roll control system to bring the missile into its proper orientation in space (flight attitude).

The roll control system operates constantly during the midcourse phase of the flight to keep the missile in the proper orientation. Roll position is obtained from the roll free gyroscope 2790 and roll rate is obtained from a roll rate gyroscope 2063. The outputs of these two gyroscopes 2790 and 2063 are mixed to give the desired amount of damping and prevent oscillations about the proper roll position.

During the time that the output signal of beam rider receiver 2042 is not utilized, the roll system can send signals to the wings 112, 113, 114, and 115 for roll stabilization as high as 150.degree. per second. After this time the tail end signals are limited to 80.degree. per second, insuring at least 70.degree. per second for roll signals. There are no separate roll control flippers but the wings 112, 113, 114, and 115 are operated differentially to cause the missile 100 to roll. Thus, no extra hydraulic system is needed for the roll system.

The guidance radar 2023 puts out its energy in the form of pulses. Each transmission will consist of the required pulses separated in time variable amounts, up to a total time for the known number of pulses of known microseconds. These will form a code group that the missile 100 can be made to receive. Any other code grouping will be rejected.

The pulse coding is necessary to prevent missiles 100 fired into different guidance beams 2024 from being confused, and for antijamming purposes.

The beam rider receiver 2042 picks up the transmissions of the guidance radar 2023 through the single teflon lens antenna 2002 that is pointed aft and is located on the missile body aft between wings 113 and 114. As the guidance transmitter 2022 (or 5412 or 5414) nutates at 30 revolutions per second, the repetition rate of the emission of the pulses is changed .+-.5 percent around the basic rate of 900 pulses per second.

The incoming pulses are fed to the superhetrodyne receiver 2042, the local oscillator 2048, of which is controlled by an automatic frequency controller referenced to the incoming TE frequency. The signal voltage produced by the receiver 2042 has a 30 cycle component, the amplitude of which is proportional to the missile error angle and its phase is proportional to the missile position in the beam. This is fed in quadrature into a pair of phase comparators 2034 and 2036. The reference voltage is obtained from the 30 cycle frequency modulation of the radar pulse repetition rate. The outputs of the phase comparators 2034 and 2036 are proportional to the error angles in each of the planes.

The missile 100 never knowns exactly how far from the center 2032 of the beam (axis of nutation of the radar) it is, since the amplitude modulation of the 30 cycle voltage is proportional to missile angular error. That is the angle, measured at the radar 2023, between the axis of nutation 2032 and the missile 100. In the computer, a motor driven linear potentiometer picks off more and more of the error signal as a function of time, thus converting the angular error to one of distance, since time can represent range in this constant velocity missile 100. However, this occurs only during the first 13 seconds after separation; after this time it is assumed that any error will be small and it can be used directly without conversion.

During this time, the guidance control system 214 is acting upon the signals received from both the roll stabilization system 2040 and the beam rider receiver 2042 and using them to position the wings 112, 113, 114, and 115. The servos in the guidance control system 214 have feedback loops to tell them just what the wings 112, 113, 114, and 115 are doing in response to the signals sent to them. In addition to these feedback loops, there is the overall aerodynamic feedback loop that measures the missile motion and feeds this back into the system to cancel commands present and executed.

The homing guidance sub-system 2014, as previously mentioned, is a semiactive radar interferometer device which produces steering signals by measuring the angular rate of change of the missile 100 to target line of sight 2052. This measurement is done in each of the planes of the wings 112, 113, 114, and 115 and so two complete systems are present, one supplying commands to the odd wings and the other the even wings. The switch over from beam riding to homing is automatically accomplished when the homing guidance sub-system 2014 receives sufficient energy from the target 2000.

The tracking radar 2028, under orders from the computer, tracks the target 2000 only during the time that the missile 100 is in its homing phase. This radar energy is reflected from the target 2000 into the four teflon homing antennae 20004, 2005, 2006, and 2007, in the nose of the missile 100. The signal from one antenna, say 2006, in channel "A", passes through a continuous microwave phase shifter 2218, and is then added to the signal from the other antenna 2004.

This produces a pulsed signal which is amplitude modulated at the frequency of the phase shifter 2218 and by the relative angular velocity of the target 2000 with respect to the missile 100 to target line of sight 2052. Thus, any deviation in the frequency of the amplitude modulation of the incoming signals from that produced by the phase shifter 2218 is due to either the motion of the target 2000 or missile 100. There are two homing rate gyroscopes 2060 and 2061 to measure missile motion and this is subtracted from the signal to give only target motion.

The automatic range tracking unit 2020 receives a signal from the beam rider receiver 2042 to start its time sweep. It produces a one-fourth microsecond gate signal which is applied to the gate video amplifier of each homing channel. Before homing guidance commences this gate sweeps in range from about 7000 feet in front of the missile 100 to about 26,000 feet.

If a signal (echo from the target 2000) appears in this gate in the odd A channel, sweeping is discontinued to the odd channel, and it is assumed that if a target 2000 appears in one channel it will also appear in the other channel. The target echo has to be present and steady, and the automatic range tracking unit 2020 has to track it for one-half second before the beam riding guidance sub-system 2012 is turned off and the homing error signal is sent to the guidance control system 214.

If, for any reason, the echo fades or is lost, there is a one-half second memory that continues to send signals to the guidance control system 214 before the homing guidance sub-system 2014 is turned off and the beam riding guidance sub-system 2012 is turned back on. (The guidance systems 2012 and 2014 are not actually turned off. Their outputs are simply switched in or out of the guidance control system 214.) There is integrated roll control during the homing phase, which limits the roll rates to less than 1.degree. per second. During this time the output of the roll free gyroscope 2790 is disregarded and the sensitivity of the roll rate gyroscope 2063 is increased.

The safety and arming device 132 is used in connection with the fuze 182 to prevent premature detonation of the warhead 194. Before detonation of the warhead 194 can be effected, this safety and arming device 132 must sustain an acceleration of at least 10 g's for 21/2 seconds, (boost) and the preset timer must run down. This time is set into the missile 100 before firing and can be either 9 or 18 seconds. When both of these functions have been met, the switch 4962 is closed completing the electrical circuit between the fuze 182 and the detonators 4800, and the metal door 4958 in the flame path from the detonators 4800 to the warhead 194 is opened, thus completing arming the warhead 194 at a safe distance from the firing ship.

When the missile 100 comes within the lethal range of the target 2000, the fuze 182 sets off the detonators 4800, which, in turn, explode the warhead 194. Fragments from the warhead 194 upon reaching the target 2000 either destroy or severely damage it.

Obviously many modifications and variations of the present invention are possible in the light of the above teachings. It is, therefore, to be understood that within the scope of the appended claims the invention may be practiced otherwise than as specifically described.

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