US3708139A - Missile control system - Google Patents

Abstract

1. A control system for steering a zero velocity launched guidable tandem coupled booster missile airborne vehicle during the boost phase of flight; comprising means for sensing the deviation of the vehicle in flight to provide steering signals, means for actuating the vehicle aerodynamic surfaces, means for supplying power to said actuating means, said power supply means being precharged before launching of the vehicle whereby power is available immediately upon the launching of the vehicle to operate said actuating means during the beginning portion of the boost phase of flight, said power being coupled to said actuating means, said sensing means comprising means for measuring the magnitude and direction of vehicle roll for providing roll stabilization steering signals, means for determining the attitude of the vehicle from the launching vector during boost phase and establishing boost phase steering signals, means for mixing the roll stabilization steering signals with the boost phase steering signals, said steering signals from said mixing means being applied to said actuating means to operate the vehicle aerodynamic surfaces thereby changing the attitude of the vehicle and the direction of vehicle flight by movement of said vehicle aerodynamic surfaces, whereby vehicle roll stabilization and vehicle attitude stability is provided throughout the boost phase of flight enabling capture of the missile in a narrow radar beam at the end of boost phase of flight.

Description

[191 r [111 3,708,139 Wh eler [4 1 Jan. 2, 1973 [54] MISSILE CONTROL SYSTEM launched guidable tandem coupled booster missile air- [75] Inventor: Phillip R. Wheeler, Alexandria, Va. 9? vehlcle during E boost of night; 3

prising means for sensing the deviation of the vehicle Asslgneei The United States of America as in flight to provide steering signals, means for actuatl'elpl'esemed by the Secretary of the ing the vehicle aerodynamic surfaces, means for sup- Navy plying power to said actuating means, said power [22] Filed: Jan. 19, 1959 supply mfans beirLg precharged befo'f launching of the vehice where y power is availa e immediately [21] Appl' 787782 upon the launching of the vehicle to operate said actuating means during the beginning portion of the [52] U.S. Cl ..244/3.13 boost phase of flight, said power being coupled to said [51] lint. (Bl H. ..B64g 1/20 actuating means, said sensing means comprising [58] Field of Search ..244/l4, 77, 79, 3.13; means for measuring the magnitude and direction of vehicle roll for providing roll stabilization steering signals, means for determining the attitude of the vehi- References Cited cle from the launching vector during boost phase and establishing boost phase steering signals, means for UNITED STATES PATENTS mixing the roll stabilization steering signals with the 2,932,467 4/1960 Scorgie ..244/77 B boost phase steering ignals aid steering ignals from 31'1'15 et a1 said mixing means being to said actuating 2,824,711 2/1958 Porter means to operate the vehicle aerodynamic Surfaces ass: 21:22; as: 1;; the of and direction of vehicle flight by movement of said vehicle aerodynamic surfaces, whereby vehicle roll stabilization and vehicle attitude stability is provided Primary Examiner-Rodney D. Bennett, Jr. Assistant Examiner-Daniel C. Kaufman Attarney-R. s. Sciascia and T. 0. Watson, Jr. throughout the boost Phase of flight enabling capture of the missile in a narrow radar beam at the end of EXEMPLARY CLAIM boost phase of flight. 1. A control system for steering a zero velocity 13 Claims, 10 Drawing Figures I 7 i Sugars WCDB wcoa SYSTEM 30 I I -"B"Plona wcos i i Output CONTROL SYSTEM 20 Wplune I fi Gyro 2 Rate Gyro COOVBFSIOII 1 i "B" Plane Unit 3 I Gyro 'sl'Y s r l i'l Amplifier 1; a 1o iIij I I ana Feed Book I I- Potentiomeler 2Q Contr =2: System 1 5 II I "B'IDC Error I I L l 3.54: f

v -J L was Gyro g 2OLimiter i2 Roll Stabilization 1 Channel Q5 l L r'irsw'mar' l y 4! LP to Sump I Potcntiomslor 2Q L Nitrogen Pressurizatlon WW 2 W SHEET 1 UP 6 PHILLIP R. WHEELER ATTORNEYS PTENTEDJAM 2197s SHEET 5 OF 6 INVENTOR ATTORNEYS PHILLIP R. WHEELER 0E6 wE 2765 S MISSILE CONTROL SYSTEM The invention described herein may be manufactured and used by or for the Government of the United States of America for governmental purposes without the payment of any royalties thereon or therefor.

The present invention relates to a missile control system for guided missiles having missile bird and booster portions, and wherein the missile combination includes wing control surfaces for providing control during free flight. The missile bird portion advantageously employs apparatus for providing wing control during the boost phase of the missile in its flight along its trajectory. This control is effective immediately after firing and is continuously active until the missile bird and booster portions separate.

More particularly, the invention relates to a system to provide stability and control of dispersion for a weapon configuration which is aerodynamically unstable through the use of aerodynamic control surfaces (wings). The motion of the aerodynamic surfaces correlates to the weapons tendency to deviate from a stabilized condition and provides corrections to the missiles attitude in order to control the weapons flight along the launching trajectory.

Prior missiles, as disclosed in copending application of Wilbur H. Goss et al, Ser. No. 594,067, filed June 26, 1956 for Guided Missile, have incorporated a larger booster having fins of considerably greater surface area and in staggered cruciform disposition along the longitudinal axis of the missile.

In previous missile control systems, guidance control started at the time of capture of the missile by the radar beam. Roll stabilization was established after separation of the bird portion and booster portion thereby exhibiting great dispersion of the missile prior to capture by having a free flight or uncontrolled portion of missile flight.

Since it is logistically desirable to utilize the same type of missile for both land based and shipboard use, any reduction in size of the overall missile is of considerable importance with respect to shipboard applications. ln order to accommodate a greater number of missiles in the stowage facilities provided on a vessel it is deemed desirable to attempt to reduce the size of the booster by use of improved solid propellant grains and thereafter direct attention to means for reducing the area ofthe fin structure of the missile. The present state ofthe missile booster art permits such an expedient.

Most current versions of shipboard handling and launching systems are of a nature incorporating semiautomatic equipment wherein the missiles are stowed in magazines on board ship and moved therefrom to ready service magazines in a semi-assembled Wingless condition with a missile bird portion and a booster portion in longitudinally joined relationship. Thereafter the individual missile to be launched is advanced or moved to a position or station where the wings and fins are assembled thereto. This occurs in an enclosed area of the deckhouse just prior to ramming onto the launching rail. As the missile is advanced from the fin assembly station in the deckhouse across a spanner rail on to the launching rail of the launcher head assembly it passes through opened blast doors which, following or simultaneously with retraction of the spanner rail, 6

The use of a large radially projecting fin assembly of the type shown in the Goss et al application supra, necessitates a large exit opening in the deckhouse and consequently a large and massive blast door. The instant invention obviates the foregoing requirement to a substantial degree by permitting use of the afore-mentioned improved shorter booster together with a reduced fin area thereon and on the missile. This facilitates stowage of a greater number of missiles, or alternatively, stowage of the same number of missiles in a smaller area. Thus the shipboard fire power capabilities may be increased and the requirements for stowage of a larger number of missiles and boosters in assem' bled relationship may be met within the existing space allocations. It further simplifies the structural requirements for handling the missile bird and booster sub-assembly in the ready service areas, the checkout areas and the fin assembly areas of the launching system. These factors all work together and, with the accomplishment of the desired reduction in mass for the design of the blast doors provided at the opening from the deckhouse, effectively provide an increase in overall operational efficiency.

The aerodynamic redesign of the foregoing type of missile for use of a booster of shorter length placed the booster fins nearer the combined center of gravity of the missile-booster combination, thus the fin moment from the center of gravity is less effective. The use of considerably smaller booster fins made the missilebooster combination statically unstable. Automatic stabilization of the aerodynamically unstable missilebooster combination as it moves along the initial and boost portion of its trajectory and the contemporaneous maintaining of stabilized flight, as a positive control within certain allowable limits of deviation from the launching vector, has been accomplished by the wing control during boost system of the instant invention wherein both the missile and booster have fixed fins. Additionally, it embodies maximum utilization of certain of the existing missile guidance components which had previously been dormant during the boost phase, namely, the roll stabilization circuits and the wing control surfaces and the associated control apparatus therefor, which were formerly utilized for effecting stability and control of the missile only during the guidance and homing phases of missile flight. The requirement for a shorter booster length and considerable smaller booster fins for compatible shipboard use has been accomplishable by the use of the instant invention.

The basic essential considerations in the presentation of the advantages of the instant invention are herewith stated in brief form to provide a better understanding of the description which follows.

The initial missile thrust for launching the missile from its zero length launching rail is obtained in the case of the instant type missile from a solid propellant booster rocket having a fitting at the forward end thereof for substantially rigid releaseable attachment to the bird portion of the missile. The missile bird propulsion system is essentially a ram-jet type of device which requires an initial acceleration to a predetermined velocity before thrust can be obtained from the ram-jet engine. This is best provided by the solid propellent grain type booster mechanism. After the booster energy has been spent, a separation occurs between the booster and missile bird portions of the missile combination. In response to this separation the control of the movable wing surfaces of the missile bird is transferred from the wing control during boost system of the instant invention to the normal guidance control system.

One object of the instant invention relates to the provision of a new and novel system for controlling within predetermined desired limits the stability and attitude of a missile with respect to the launching vector therefor, during the portion of missile flight wherein the missile and booster combination are traveling as a unitary device with both missile portion and the booster portion having fixed fins and with the boost phase of the missile trajectory being controlled by missile wing movement.

Another object of the present invention is to provide improved stability during the boost phase of missile flight without the necessity for usage of large fins on the booster.

Another object of this invention resides in the provision of instrumentalities to inhibit large initial deviations of the missile from the launching angle or vector so that the missile will be in the most opportune position for capture by the radar beam and subsequent beam riding control after transfer of such control functions thereto at the completion of the separation sequence of the missile in flight.

In correlation with the preceding object a further object of the present invention is to enhance the subsequent stabilization during the boost phase by the same and certain additional instrumentalities by minimizing or correcting deviations due to aerodynamic and thrust misalignments and thereby better enable capture of the missile in a narrow guidance beam after separation of the missile from the booster.

Still another object is to overcome adverse influences upon the missile launching attitude which might result from cross winds and/or boost fin angular misalignment.

In correlation with the immediately foregoing object it is a further object to incorporate in the sensing equipment, features providing for anticipation of tendencies to deviate from the desired missile attitude and to effect correcting influences or control over the missile thereafter.

Another object of the present invention is to provide in combination with a wing control during boost system, fixed fins on the booster of a size substantially reduced from prior booster fin configurations, which are disposed in line with the missile wings and fins to present a symmetrical disposition about the missile air frame axis, thereby to add to the overall stabilization of Another object of the present invention is to enable capture of the missile by the guidance radar beam at the earliest possible time by the incorporation of roll stabilization during boost phase of missile flight.

Another object of the instant invention resides in the provision of instrumentalities for control of the missile wings during the boost phase of missile flight along its trajectory by utilizing to a maximum extent the existing control mechanisms incorporated in the guidance control system of the missile and with the utilization of a minimum of additional weight adding apparatus thereto.

Still another object of the instant invention resides in the provision of structure and apparatus for overcoming all of the afore-mentioned shortcomings of missile and booster systems heretofore or now in general use while providing substantially all of the desirable features and functions thereof without sacrifice or compromise of performance during the boost phase of operation.

In correlation with the immediately foregoing object it is further object to accomplish the foregoing desired performance characteristics without sacrificing performance of the missile during the terminal phases of operation when the missile is under control of the guidance control system.

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 generally diagrammatic illustration, with the missile outline portionthereof in perspective form, to indicate certain of the physical relationships of the instant invention;

FIG. 2 is a longitudinally exploded perspective view showing the electrical system package and the hydraulic control system in their respective positions in an assembled relationship in a segmental portion of the missile bird;

FIG. 3 is a diagrammatic illustration in block form showing the electrical, hydraulic and mechanical mode of function of the invention;

FIG. 4 is a diagrammatic illustration of the hydraulic system utilized in the missile for use with the instant invention;

FIGS. 50, 5b and 5c are operational schematic diagrams of the servo valves and wing actuators of the hydraulic system of FIG. 4;

FIG. 6 is a schematic diagram of the electronic system of the rate conversion unit of FIG. 3;

FIG. 7 is a schematic diagram of the electronic system of the attitude stabilization channel of FIG. 3; and

FIG. 8 is a block diagram of the roll stabilization system. 7

Referring now to the drawings wherein like reference characters are utilized to designate like or corresponding parts throughout the several views and more particularly to FIG. I wherein there is shown by way of example in a somewhat diagrammatic form the outline of a surface to air type of a supersonic missile. The showing is of a two stage type guided missile in which the first stage or booster is of the solid propellent rocket variety for launching and accelerating the missile to a predetermined velocity. The second stage is a ram-jet propelled missile bird portion wherein the ram-jet propulsion apparatus is rendered active when the missile has been accelerated to sufficient velocity for ramjet action to take place.

The missile generally indicated at is provided with four cruciform-arranged, biconvex movable wings of low aspect ratio and having a raked-tipped configuration. In this arrangement these low drag pairs of movable wings I, 3 and 2, 4 respectively, are disposed approximately at the missile bird longitudinal center of gravity. The two respective oppositely disposed pairs of wings, which are arranged in the afore-mentioned cruciform configuration, define reference planes of the missile hereinafter referred to as the A" plane and B plane.

The air frame portion of the missile bird which is shown merely in outline form in FIG. 1 also includes four rectangular-planform fins of biconvex cross-sectional configuration which are indicated at and located near the aft section thereof. It is to be noted that both the fins 15 and the movable wings I to 4 are disposed in a longitudinally aligned relationship along the missile bird air frame.

The aft end of the booster portion of the missile, from which the initial missile thrust is obtained from the solid propellant disposed therein, supports four single spar-type fins 17 of biconvex construction and trapezoidal planform configuration. Additionally the fins 17 have a tapered leading edge.

A better appreciation of the relative magnitude of the reduction in the size of the booster fins as utilized with the missile of the instant invention as related to the extremely large booster fins of a prior type missile of Goss et al, supra, as shown at 18 on FIG. l in broken outline representation, will be apparent by a comparison with a typical booster fin as shown in heavy outline at 17 which is representative of the fin for the missile booster of the instant invention.

The prior discussions relating to the flight of the missile from the time it leaves the zero length launching rail to the intercept with the target, or to a target destructive proximity in the target area, may be considered as being divided into three phases, wherein the launching'and-boost phase herein referred to as the boost phase is the first or initial portion, and the one with which the instant invention is primarily concerned. The second portion is a separation phase wherein a physical separation between the spent booster and the bird portion of the missile occurs along the trajectory thereof. The last or terminal portion of missile flight is the beam riding or guided flight phase.

In the event that some sort of terminal homing facilities are provided in the missile intelligence system this portion of missile travel would be considered as being included generally in the guided flight phase.

Tactically the guided missile may be considered as being launched from a zero-length shipboard launcher of either a single or a dual rail type which is capable of being trained in azimuth and moved in elevation to a predetermined launching or firing angle. It is to be understood however that the same general relationships will apply in the event the missile is launched from a land based launcher arm or launcher cart which is likewise of a character providing movement in train and elevation. This firing angle corresponds to the angle of the launching vector. The instant invention is primarily concerned with the establishment and the continuous maintenance of coincidence with the vector during the boost phase of missile flight.

The positioning of the launcher prior to missile firing is preferably determined by a control system using servo systems for relaying electrical or hydraulic signal intelligence from a fire control system to the servo follow-up system for train and elevation of the launcher head carrying the launcher cart or launching rail as the case may be. During the readying period before the missile is launched but while the missile is retained on the launcher, the control system electronic packages as enclosed in the missile are warmed up from an external source of electrical power for the purpose of carrying out semi-automatic prelaunch checkouts and for providing settings of the necessary controls with the launcher. Additionally indications are communicated from the missile to the control system indicating the condition of the various controls. Upon completion of the loading check-out, and somewhat prior to the launching of the missile, the control system indicated generally at 20, FIGS. 1 and 2, of the missile is switched to the missile internal power system. The boost phase may be considered to begin when the booster rocket is ignited by an electrical command. As will become apparent as the description proceeds certain of the reference controls are obtained by use of gyro type mechanisms which are started during the period immediately prior to launching whereby they may be accelerated to operating speed to obtain sufficient momentum to provide the desired controlling functions after the missile is accelerated from the launcher. Upon ignition of the booster rocket 16 the missile l0-andbooster l6 combination is accelerated from the zero length launcher to obtain the necessary speed required for ram-jet operation.

During the boost phase movement of the missile, wings 1, 2, 3 and 4 disposed in the A" and B planes function in conjunction with the stabilization control provided by the fixed fins on the missile and on the booster and indicated respectively at 15 and 17 to provide correction for any deviation of the missile attitude from the launcher aiming vector.

The device or apparatus for sensing deviation of missile attitude from the launching vector during the boost phase in missile flight and for initiating command signals which are ultimately applied to the wing servo system, comprises the principal subject matter of the present invention. This system will be referred to as the wing control during boost system and is hereinafter designated as WCDB.

The WCDB system 30 as hereinafter described in greater detail, is advantageously so incorporated into a missile control system 20 as to obtain a maximum utilization of existing system component elements with a minimum of additional function and control equipment. The overall WCDB system is functionally indicated in the block diagram of FIG. 3.

The instant system is rendered active at launching of the missile and functions during the initial or boost phase to provide full control over missile flight. In this action the boost control system 30 derives error or deviation signals relative to any deviation of the missile flight from the desired launching vector and communicates the essential boost correction command signals from the gyro system through relay contacts 70 to the mixer amplifier 71 wherein the boost signals are combined with the roll stabilization signals which are derived in a similar manner hereinafter described in greater detail from the roll stabilization channel 63. From the mixer amplifier 71 the signals are demodulated to provide DC or low frequency AC signals which are applied to the input stage of the respective servo amplifiers 73, there being a corresponding servo amplifier for each of the four wings 1 to 4, respectively and substantially as shown. The respective servo amplifier for control of each wing surface applies the signal to the input of a DC differential amplifier which in turn is followed by a push-pull power output stage. The amplified power output of the push-pull stage is fed through the windings of a torque motor which, in turn, functions to regulate oil flow in the hydraulic system and provide wing movement. The manner of accomplishing energization of the hydraulic system for proper hydraulic fluid flow to cause the actuation of the piston and the manner of effecting the actuating of the individual bell crank linkages for the respective wings will become more apparent from the description of FIG. 4 for the hydraulic servo system. Rotation of the respective wing surface about its individual axis of rotation functions to change the instantaneous missile flight attitude by an amount necessary to correct for any initial missile flight deviation occurring during the boost phase. The follow-up control, hereinafter described in greater detail is included in the system to maintain the missile attitude within desired limits by sensing deviation tendencies of the missile and thereafter applying the necessary control as an electrical feedback signal as circumstances of the occasion may necessitate.

At the completion of approximately 5 seconds of missile-booster flight, the booster rocket thrust diminishes and aerodynamic drag and diffuser pressure cause the booster 16 to separate from the bird portion of the missile and fall free therefrom. At the instant of separation, separation microswitches 19 FIG. 1 which are normally disposed in open circuit relation ship by the physical contact of portions of the booster which project to engage the actuators therefore in a conventional manner cause circuit closure through the microswitch to provide energization of the primary circuit of the separation relay in the electrical system. When suitably energized this relay functions to connect certain of the missiles internal battery power sources in a manner to apply a voltage to the ram-jet ignition unit and to provide an electrical separation command signal for the missile control system. The application of the separation command signal to the missile control system functions to actuate relay contacts 70 thereby disconnecting the WCDB system 30 from the mixer amplifier 71 and connecting a guidance channel control system 51 therewith for providing subsequent control over the wing surfaces by means of the guidance control.

The guidance control is established by guidance signals which indicate error in the flight of the missile from the guidance transmitter radar beam. The guidance beam of the radar is positioned during launch to permit the beam to capture the missile following separation and is programmed during the remainder of flight in a manner required to steer this missile to the immediate vicinity of a target. The necessary intelligence is transmitted to the missile by means of a pulse-coded nu'tated and pulse-repition-rate frequency modulated microwave signal. in the control system 20, a beam rider receiver in the intelligence system 50 connected to an aft mounted beam rider antenna 21, de-- tects the intelligence in the nutation of the beam and the frequency modulation of the pulse-repetitionfrequency required to steer the missile to the center of the guidance transmitter beam and maintain that position throughout the missile flight. The guidance control system 51 receives from the intelligence system 50 through the multiplexing switch 52 two signals corresponding to the error in the A and B planes of the missile and modifies these signals according to missile altitude as determined by the static pressure switches 53 to obtain electronic commands to effect proper movements of the missile wings. The modified signals from the guidance control system are multiplexed and then fed to the mixer amplifiers 71 through the relay 70.

Before proceeding with a detailed description of the electrical boost-portion-of-flight controls for wing control prior to guidance, a description of the function and the desired mode of operation of the hydraulic system is deemed to be in order to provide a clearer understanding of the manner in which the boost phase electrical and hydraulic system elements cooperate in their respective functions to provide the power actuation for the proper wing control motion. As aforementioned the input to the hydraulic system is in the form of electrical command signals from one or more of the servo amplifier assemblies 73. These servo amplifiers 73 advantageously provide a power output sufficient to electrically energize the respective torque motor 91 for its pilot valve 93 on the respective hydraulic control valve 74 of the servo system 89 connected thereto. The wing servo valves 74 provide suitable porting of fluid under pressure for actuation of the differentially operable pistons or actuators 78 which are connected by means of bell crank linkages 4 for rotation of the wing surfaces 1 to 4 about their respective axes. These axes are of an orientation radial to the longitudinal axis of the missile and lie in a common plane.

This hydraulic power portion of the system generally includes a hydraulic power supply circuit having therein an air turbine driven hydraulic pump 81, and a suitable check valve 82 interposed in the output line therefrom. The check valve 82 functions initially to prevent bleed off of precharge pressure in the accumulator circuit or path 92 of the system, and thereafter permits take over of the charging function for the system to permit buildup of sufficient pressure in the accumulators 84 after pump 81 energization during boost phase missile flight to provide continuous hydraulic system control over guidance phase flight.

A solenoid controlled valve 83 provides the means whereby a suitable electrical current flow initiates operation of the hydraulic system 79 for high pressure fluid flow through the servo valves 74 which provides fluid flow of predetermined alternatable direction for the actuating pistons of the wing actuators 78. The system additionally includes the aforementioned high pressure accumulators 84 which are precharged with nitrogen to provide sufficient initial pressure for hydraulic actuation of the wing control surfaces 1 to 4 during the beginning of the boost phase of missile operation and/or at least prior to air turbine actuation of a magnitude sufficient to drive the air turbine driven hydraulic pump 81. The feed from the high pressure hydraulic accumulator 84 is through the main manifold system 85 which includes hydraulic pressure switch 86, pressure relief valve 87 and a low pressure sump 88. The individual elements of the hydraulic system may be considered as providing functions corresponding to a conventional hydraulic system of similar elements for the purposes mentioned. The servo control portion of the hydraulic system as indicated generally at 89, FIG. 4, includes the four serve valves 74 and the four actuating pistons 77 of actuators 78 as well as the feedback potentiometers 90 connected to follow the bell crank linkage movement 94 and feedback an electrical signal for electrical signal reference purposes.

In operation of the system, suitable hydraulic fluid, hereinafter referred to as oil and contained in the low pressure sump 88 is pressurized by the hydraulic pump 81 and circulated through a port past check valve 82 through a high pressure line or path 92 to the accumulators 84. This hydraulic system is a closed circuit arrangement wherein the precharged accumulators feed the wing servo system 89 through the path 92 to the main manifold 85 to the separate servo valves 74. As aforementioned electrical signals from servo amplifier 73 function to regulate the flow of oil to and from the wing actuators 78. Oil flow is of a reversible unidirectional character for providing what might be termed up-and-down movement of the wing control surfaces 1 to 4. Oil from the low pressure side of pistons 77 of actuators 78 is returned through the return ports of the servo valves 74 back to the sump 88 to complete the flow circuit of the closed hydraulic servo loop. When the wing actuators 78 and wings move, electrical signal intelligence in the nature ofa resistance variation across portions of the feedback potentiometers 90 is applied to the servo amplifiers 73 in a manner which will become more apparent as the description proceeds. This action provides a positive indication of the missile wing position and completes the feedback circuit to the servo amplifiers 73. Functionally, the hydraulic system is divided into the following two separate systems: the hydraulic power supply system 80 and the wing servo system 89.

The hydraulic power supply system 80 comprises a hydraulic pump 81, an air turbine 95, FIG. 2, a turbine power control generally indicated at 96, two high-pressure accumulators 84, a high-pressure manifold 97, a main manifold 85, and a low-pressure sump 88. The system check valve 82 and a solenoid-operated shutoff valve 83 are located at the high-pressure manifold. The system filters 98, a hydraulic pressure switch 86, and a pressure relief valve 87 are mounted on the main manifold 88 as indicated on FIG. 4.

The hydraulic system must react rapidly to input guidance signals and must position the wings accurately in relation to the signals. To meet these requirements, the system must have certain basic components designed to function properly in relation to each other in the overall system.

The wing servo system consists of four servo valve and wing actuator combinations. Two servo valves are mounted on the manifold block in quadrant 1, FIG. 2, and two servo valves are mounted on the manifold block in quadrant 3. Each actuator 78 is connected mechanically to a missile wing and its respective feedback potentiometer 90.

The hydraulic system is designed to meet the maximum power conditions expected at the wings during missile flight. The two factors controlling the hydraulic power requirements are: (1 the torque needed to turn the wings, and (2) the angular velocity at which the wings move. These two factors determine the size of the wing actuator piston and cylinder and the radius of the bellcrank attached to the wing sleeve 101.

For a fixed wing actuator and bellcrank combination, the maximum hydraulic power requirements at the wings can be considered in terms of an oil pressure differential across the actuator piston and the total oil flow rate to and from the actuator cylinders. The power requirements of the hydraulic pump, however, must be based on the pressure rise required between the sump and the accumulators and the total oil flow rate in the system.

The oil pressure applied to the actuator piston produces a force on the end of the bellcrank attached to the wing sleeve 101. The torque applied to the wing sleeve 101 is the product of the piston force and the effective bellcrank moment arm. This torque must balance the maximum wing hinge moment expected during missile flight.

For a constant missile flight speed, the wing hinge moment is dependent upon the altitude and the wing incidence angle. Hinge moment increases with an increase in incidence angle and decreases with altitude. During flight, at altitudes under 35,000 feet, the turn radius of the missile must be limited to permit no more than a 15 g turn. This requires limiting the wing positions to values under the 20 maximum incidence angle. Above 35,000 feet, the missile cannot exceed a 15 g turn even with the wings at the maximum 20 angle. The maximum wing hinge moment is limited to 9400 inch-pounds up to approximately 35,000 feet. Above this altitude, however, the hinge moment drops off and is only 3500 inch-pounds at 60,000 feet.

The torque applied to the wing is a product of the hydraulic oil pressure differential across the actuator piston, the effective piston area, and effective moment arm. The effective moment arm varies between a minimum of 2.82 inches and a maximum of 2.99 inches during full wing movement. The maximum value is equal to the bellcrank radius. The actuator piston has an effective area of 2.66 square inches at its open end and an effective area of 2.51 square inches at the rod end, giving an average effective area of 2.58 square inches.

On the basis of the average effective piston area and the minimum moment arm, a pressure differential of 1250 p.s.i. is required at the wing actuator to balance the maximum wing hinge moment (9400 inch-pounds).

The oil flow required by each actuator is detennined by the nominal wing angular velocity. This velocity is per second. Wing angular velocity corresponds to the rate of movement of the piston in the actuator cylinder and therefore determines the rate of hydraulic valve pressure drop, together with the actuator load pressure results in a total system pressure drop of l850 p.s.i.

The total power requirement for the overall hydraulic system is based on the total system pressure drop (1850 p.s.i.) and the total oil flow requirement. The total oil flow requirement is a total of the oil flow required for all four wing actuators and all four servo valves, or 80.5 cubic inches per second. This results in a total power requirement of approximately 22 horsepower.

As hereinbefore setforth the major components of the hydraulic power supply system are the hydraulic pump, air turbine, turbine power control, high-pressure accumulators, and low-pressure sump. Minor components of the system include the system check valve 82, solenoid-operated shutoff valve 83, filters 98, hydraulic pressure switch 86 and pressure relief valve 87.

The hydraulic pump circulates the oil from the lowpressure sump 88 to the high-pressure accumulators 04 and supplies the oil at a constant pressure differential to the wing servo system.

The use of the pump in the hydraulic system eliminates the need of packaging a large supply of pressurized oil in the missile. This reduces the missile weight and makes space available for other missile components. Due to the relatively low-flow and highpressure characteristics of the system, a constant displacement-type pump is used. The pump 81 is an axialpiston type of a conventional nature containing seven pistons in a rotating cylinder or rotor. The pistons reciprocate as the cylinder rotates due to the action of an angular cam plate.

The pump output must be equal to the maximum oil flow demands of the system at a system pressure rise of 1850 p.s.i., which as aforestated is equivalent to approximately 22 horsepower. Since the pump has an efficiency of 80 to 85 percent, at maximum oil flow and pressure rise conditions, an input to the pump of approximately 27 horsepower is required. The power input to the hydraulic pump is supplied by the air turbine 95 and is transmitted to the pump through the speed reduction gears indicated generally at 99 on FIG. 4. The gears preferably have a speed reduction ratio of 6.2 to 1. While the pump rotor is turning at 3750 r.p.m., the turbine wheel is rotating at 23,250 r.p.m.

While the missile is moving in a straight and level flight, there is essentially no wing motion, resulting in a condition of nearly zero oil flow. During the no-flow condition, the constant-missile flight speed, the power available for driving the turbine decreases with an increase in altitude.

At the pump speed required for maximum oil flow demands, the turbine has an efficiency of approximate- 1y 50 to 55 percent. On the basis of the turbine and pump efficiencies, an overall efficiency of approximately 40 to 45 percent is obtained from the air-turbine-driven pump assembly. The air horse-power requirements are limited up to 35,000 feet altitude due to wing hinge moment limitations imposed on the mi sile by the control system.

The turbine power control 96 is incorporated on the air-turbine-driven pump assembly for the purpose of maintaining a constant oil pressure rise from the pump. The pressure rise is maintained at 1850 psi. above sump pressure.

As the missile altitude increases, the control progressively allows more diffuser air to the turbine to compensate for the decrease in air density. During changes in oil flow demand by the wing servo system (at constant altitude), the control acts to increase or decrease the air flow so as to correspondingly increase or decrease the turbine power. Air flow to the turbineis regulated by a throttle plate not shown located in front of the turbine nozzle plate, also not shown. At sea level the total nozzle area required is approximately 0.4 square inch at 60,000 feet, however, the nozzle must be full open, giving a total area of 3.5 square inches.

Two high-pressure accumulators 84 are used in the hydraulic power supply system. The accumulators are connected through the solenoid-operated shutoff valve 83 to the main manifold 85 and the hydraulic pump 81, as shown in FIG. 4. The accumulators supply high-pressure hydraulic oil to the wing servo system during the missile boost phase. During missile flight, the accumulators contain a reserve volume of high-pressure oil and act to smooth out the oil flow and pressure fluctuations in the hydraulic power supply system. Each accumulator contains a free-moving piston, not shown, suspended between the hydraulic oil and compressed nitrogen.

When hydraulic oil from the pump flows into the accumulators, the piston moves, compressing the nitrogen until the nitrogen pressure balances the oil pressure. When the hydraulic power supply system pressure drops, due to sudden oil flow demands by the wing servo system, the compressed nitrogen causes oil to flow into the system. This action maintains the system pressure during the time delay required for the hydraulic pump to come up to full operating speed.

A secondary action of the accumulators occurs during periods of peak oil flow, while the missile is flying at I high altitudes. During periods of peak oil flow to the wing servo system, oil must be supplied by both the pump and the accumulators. Due to the fall-off in air density at high altitudes, the air horsepower input to the hydraulic pump air turbine becomes extremely low. When the input power decrease is accompanied by peak periods of oil flow, a continuing decrease in oil pressure occurs. The system oil pressure decrease at higher altitudes is allowable, however, because of the corresponding decrease in the wing hinge moment.

The actual conditions under which oil is received by the accumulators and supplied to the hydraulic system is controlled by the initial nitrogen pressure and volume and, to some extent, by the behavior of the nitrogen during compression and expansion. The compression or expansion of the nitrogen behaves basically according to the adiabatic pressure-volume relationship (PV-=C), during which no heat is lost during compression, and no heat is added during the expansion. Since some heat transfer actually does occur, due to the metal accumulator body, the actual relationship is closer to PV C.

Nitrogen under pressure is placed in each accumulator before the hydraulic system is filled with oil. The nitrogen pressure in accumulator 1 is 1700 p.s.i.a., and the nitrogen pressure in accumulator 2 is 1 W p.s.i.a. The accumulator pistons are bottomed at the hydraulic oil end of the cylinders, giving a nominal nitrogen volume of 200 cubic inches in each accumulator. After the accumulators are charged with nitrogen, the system is filled and bled. The bleed valves utilized for this purpose are indicated by ledgends on the various components as indicated on FIG. Low-pressure oil, at 30 to 40 pounds per square inch g age, is retained in the system. The pistons remain in the bottomed position, since the nitrogen pressure is greater than the oil pressure.

Before the missile is launched, oil at 2500 p.s.i.a. is introduced into the accumulators at the high pressure oil fill port. This pressure is retained in the high pressure manifold and accumulators 84 by the normally closed solenoid-operated shutoff valve 83. As the quantity of oil increases the accumulator pistons move, compressing the nitrogen until the gas pressure balances the oil pressure of 2500 p.s.i.a. Since the initial nitrogen pressures differed in each accumulator, the positions of the pistons will differ at the balanced condition. With the accumulators charged with oil at 2500 p.s.i.a., approximately 64 cubic inches of oil is stored in accumulator 1, and approximately 112 cubic inches of oil is stored in accumulator 2. The pressurized hydraulic oil is stored in the accumulators for use during the missile boost phase.

In prior art systems not incorporating the wing control during boost phase feature of the instant invention, the accumulators of typical systems did not function to supply any oil under high pressure until such time as the missile had accelerated to sufficient speed to actuate the air turbine driven hydraulic pump for accumulator charging. Since in the instant system oil under high pressure is needed immediately upon launching of the missile to provide the necessary power for operation of the hydraulic servo system during the initial portion of missile boost, hydraulic oil in the accumulators is precharged by nitrogen charging to a required pressure. This nitrogen charged hydraulic oil in the accu mulators is initially sealed off at the outlet of the highpressure manifold block for the accumulators by the solenoid-controlled shut off valve 83. At approximately two seconds before missile firing, solenoid valve 83 is energized, thereby releasing the pressurized oil for flow from the accumulators 84 to the hydraulic manifold 85, thence to the four servo valves 74. During the first two seconds of the boost phase the hydraulic pump is not operating and oil for the wing control actuators flows from the accumulators 84.

At approximately three seconds after boost, hydraulic pump pressure build up equals the oil pressure in the accumulators and pump 81 starts to replenish pressure in the accumulators. During missile flight, the pressure in the accumulators is maintained at approximately 1900 p.s.i.a. The actual pressure is 1850 p.s.i. above sump pressure and depends on the pressure rise in the low-pressure sump during the first two minutes of the boost phase.

During the missile flight, one accumulator acts primarily to smooth out pressure fluctuations, and the other accumulator acts primarily to store a volume of high-pressure oil which smooths out volume fluctuations. The difference in performance between the two accumulators is a result of the initial nitrogen precharge pressures of 1700 p.s.i.a. and 1100 p.s.i.a.

The low-pressure sump is located on the servo valve manifold in quadrant 1, as shown in FIG. 2. The sump receives oil returning from the wing actuators of the wing servo system and acts as an oil supply for the hydraulic pump. A bladder in the sump containing compressed nitrogen pressurizes the oil at the pump inlet port, thus preventing oil cavitation. Secondary functions of the sump include storing oil to reduce temperature buildup and supplying makeup volume for any system leakage.

Before the sump is charged with nitrogen, atmospheric pressure is present in both the cylinder and the bladder. With the hydraulic oil end of the cylinder open to atmosphere, the bladder is charged with nitrogen to 20 p.s.i.a. With the sump in the nitrogen charged condition the bladder expands in the cylinder and bottoms against the end cap. In this condition, the sump is mounted against the servo valve manifold in quadrant ll. in the charged condition, there is a maximum nitrogen volume of 237 cubic inches at a pressure of 20 p.s.i.a.

When the hydraulic system is filled and bled, oil at 30 to 40 p.s.i.g. is retained in the system, and the bladder is compressed until the nitrogen pressure balances the oil pressure. The minimum allowable oil pressure in the sump 38 when the missile is launched is 30 p.s.i.a. Under this condition, approximately 60 cubic inches of oil is stored in the sump.

During the initial two seconds of the missile boost phase, hydraulic oil is expended to the sump 88 from the wing actuators 78. Since the hydraulic pump does not start to circulate oil until approximately two seconds after launch, oil is not being returned to the high-pressure accumulators. A pressure buildup in the sump to approximately p.s.i.a. could occur under maximum anticipated oil flow demands to the wing actuators 73. a portion of the oil volume will be removed and the pressure will be reduced. Under normal missile flight conditions; the sump pressure will be reduced to a nominal pressure of approximately 50 p.s.i.a., and approximately 1 15 cubic inches of oil will be stored. During missile flight, the nitrogen pressure in the bladder maintains a back pressure on the inlet port of the hydraulic pump, preventing cavitation. The volume of oil stored in the sump provides an oil reservoir for the pump and, at the same time, acts to slow down the oil temperature rise in the hydraulic system and provides a makeup volume for system leakage.

During missile flight at high altitudes, it is possible for the accumulators to discharge oil to the wing servo system @9 and the sump faster than the oil can be returned to the accumulators. This is due to the fall-off in available air horsepower to operate the hydraulic pump. To prevent damage in the low-pressure side of the hydraulic system, the sump pressure rise is limited to 120 p.s.i. by the pressure relief valve 37 mounted on the main manifold block 85, FIG. 4. The relief valve dumps oil overboard if it is necessary to relieve the sump pressure.

During the boost phase, special stabilization is required in order to hold the missile-booster combina tion to a path parallel to the launcher aiming vector. The stabilization minimizes deviation caused by wind or due to aerodynamic and thrust misalignments. This enables capture of the missile in a narrow guidance beam at separation of the missile from the booster. Stabilization is accomplished by the movement of the wings during the boost phase. The accumulators, which are precharged with pressurized hydraulic oil before the missile is launched, supply the hydraulic power required by the wing servo system during the initial portion of the missile boost phase.

The hydraulic system diagram, FIG. 4, illustrates the connections of the hydraulic power supply system to the electrical system, i.e., one set of connections at 102 to the solenoid valve 83 and the second set of connection 103 to the hydraulic pressure switch. These are in addition to the connections at 104 to the torque motors. Prior to the missile being launched, the hydraulic oil at 2500 p.s.i.a. is sealed off in the accumulators by the solenoid valve. Hydraulic oil in the remainder of the system remains at sump pressure. At approximately two seconds before launch, external power from a power transfer relay, not shown, energizes the solenoid valve 83, releasing 2500 p.s.i.a. oil pressure to the main manifold and the four servo valves. Oil pressure to the main manifold closes the hydraulic pressure switch 86, completing the missile battery voltage or internal power circuit to a separation relay, not shown, and the activation relay 211. The pressure switch 86 will not close unless the hydraulic pressure is 1400 p.s.i. or greater. At boost, battery voltage through the pressure switch is by-passed through the G-switch, hereinafter setforth in the description of the electrical system.

Although the details of the solenoid valve 83 form no part of the instant invention this valve which is indicated in diagrammatic form only, is of a character incorporating a poppet type valve and coaxially disposed a solenoid type pilot valve therefor wherein high-pressure oil from the accumulators is transmitted through a small port located at the side of the poppet on the downstream side of the poppet and pilot valve. The imbalance of pressure acts to hold the poppet on its seat in the valve body and, at the same time, acts to hold the pilot valve against the poppet. The pressure in the accumulators at this time is 2500 p.s.i.a., while the pressure in the main manifold, and at the four servo valves, is 30 to 40 p.s.i.

Due to the large pressure against the poppet, a corresponding large force would be required to pull the poppet off the axially disposed port in the valve body. This would require a very large solenoid plunger and coil. The solenoid, therefore, is of a character designed to pull the pilot valve off the small port at the center of the poppet, a smaller force being required because of .the smaller port area involved.

When the solenoid coil at 83 is energized, the solenoid plunger 105 is pulled into the coil. This results in moving the pilot valve in the poppet, uncovering the small port at the center of the poppet, and closing off llt:

the small high-pressure oil port at the side of the poppet. Oil on the solenoid side of the poppet is thus reduced to main manifold pressure, causing an unbalanced force on the poppet. The force now, however, is away from the poppet port and the poppet is moved to its fully open position. Accumulator oil pressure is then transmitted to the main manifold and the four servo valves.

When the solenoid plunger 1105 moves into the coil, a spring-loaded locking pin (not shown) drops in place in front of one of a plurality of longitudinal lands provided on the plunger outside diameter. This prevents the plunger from returning to the released position and holds the valve off its seat. The valve thus remains open during missile flight.

There are two operating positions of the conventional type pressure switch 8 6. Oil pressure acts against the metal diaphragm therein, not shown, which is in contact with a switch contact actuator on plunger. Before the missile is launched, the pressure is not great enough to flex the diaphragm sufficiently to move the plunger into contact with the switch. The switch contacts remain in the normally open position. Just before boost, high-pressure oil from the accumulators is released to the main manifold and the wing servo system. if the pressure is 1400 p.s.i. or greater, the force exerted against the metal diaphragm is sufficient toflex the plate enough to provide closure of the switch contacts. This completes the circuit from the missile battery to the activation relay 211 in the missile prelaunch condition of the electrical system. If the pressure is below 1400 p.s.i., the circuit is not completed and the missile will not be launched.

The accumulators must supply the total oil flow requirements of the wing servo system during the first two seconds of missile boost, since the pump is not operating at this time. As oil is expended through the servo valves, the accumulator pressure is continuously dropping and is accompanied by a buildup in sump pressure. As aforementioned the hydraulic pump 81 does not start to function until the air pressure in the diffuser of the missile is sufficient to operate the pump air turbine 95. Diffuser pressure increases with missile speed and, at one second after launch, is sufficient to start pressurizing the oil on the outlet side 106 of the hydraulic pump. No oil flow from the pump is possible, however, until pump pressure buildup equals the accumulator pressure, at which time the hydraulic system check valve 82 opens. Pump pressure rise meets accumulator pressure fall-off at approximately 1700 p.s.i. After the inertia of the pump rotating parts is overcome, the pump starts to replenish the oil in the accumulators and builds up the accumulator pressure. The hydraulic pump maintains the system pressure differential at the designed value of 1850 p.s.i. above sump pressure. The sump pressure buildup depends upon the amount of oil expended through the servo valves before the hydraulic pump starts to recirculate oil in the system.

A simplified schematic of the servo valve in three different conditions of operation is illustrated in FIGS. 5a, 5b and 5c. The servo valves are regulated by electrical signals originating in the control system of the missile as hereinafter described in greater detail. The magnitude and sign of the signals are the result of guidance signals fed into the mixing amplifiers of the control system. The resulting input to each servo valve is a differential current applied to the torque motor 91.

High-pressure oil from the hydraulic power supply (approximately 1850 psi. above sump pressure) enters the servo valve at the high-pressure oil port 107 and flows around the outside diameter of the center land 108 of the valve spool 109. At this point, the oil flow is divided between two passages 111 and 112, each containing a restrictive orifice 113. Each passage is connected to a metering nozzle 115 and chamber 114 at the end of the valve spool 109. The restrictive orifices 113 meter the oil, causing the pressure to drop to approximately 500 psi. Oil at reduced pressure is then ported to the chambers 114 at the ends of the spool 109 and to the metering nozzles. The oil flows through the metering holes of the nozzles, again dropping in pressure, and flows through the drain passage 116 to the drain port 117. The drain port 117 of the servo valve 74 is connected to the low pressure sump. Continuous oil flow through the nozzles is approximately 0.1 gallon per minute per each servo valve. If the differential current in the torque motor coil is zero, the control arm 118 (reed) remains in its centered position. When the control arm is centered between the metering nozzles 115, the restriction of oil flow from each nozzle is equal, and the oil pressures at the ends of the spool 109 are also equal. With equal pressures at each end of the spool, the spool remains centered over the high-pressure oil port 107 and the two drain ports, preventing oil flow to and from the wing actuators. The centered position of the spool is illustrated in FIG. b.

When a differential current is applied to the torque motor coil at 104, magnetic forces are applied to the control arm 118, causing the arm to be deflected away from the centered position between the metering nozzles. If a differential current is applied to the coil, causing the control arm to be deflected to the left, FIG. 5a, the restriction of oil flow from the left nozzle 115 is increased, while the restriction to oil flow from the right nozzle is decreased. Consequently, the oil pressure is increased at the left end of the valve spool 109 and decreased at the right end of the valve spool. The pressure differential across the spool moves the spool to the right, connecting high-pressure oil to the left actuator port 119 and, at the same time, connecting the right actuator port 120 to the drain port and the low-pressure sump.

If the polarity of the differential current is reversed, so that the control arm is deflected to the right, FIG. 50, instead of to the left, as previously described, the oil pressure differential across the spool 109 will move the spool to the left. The porting to the wing actuator is thereby reversed.

For a constant-pressure differential in the hydraulic system, the servo valve will provide a linear relationship between the amount of differential current applied and the oil flow rate to and from the wing actuator. The above linear relationship is independent of hydraulic pressure drop through the valve as a result of the internal design of the valve. The valve gain (change in oil flow rate per change in current differential) has a nominal value of 3.17 cubic inches per second for each milliampere change in current differential. The current differential is limited to 6.75 milliamperes to limit the maximum oil flow rate to and from the wing actuator. The maximum oil flow corresponds to a maximum wing rate of 150 degrees per second (nominal value).

The design of the servo valve 74 allows a power amplification of approximately 1 to 70,000. Roughly 40 milliwatts of electrical power at the torque motor controls 3.73 horsepower (roughly 2800 watts) to the wing actuator. This amplification is accomplished in two stages. The first stage consists of the torque motor and metering nozzles; the second stage consists of the transfer valve and the valve spool.

Movement of the piston 77 in the actuator cylinder 78 is dependent upon the rate of hydraulic oil flow into and out of the cylinder chambers. The oil flow rate and direction of flow to each actuator is determined in the missile control system and is regulated at the servo valve.

The forces acting on the actuator piston are a result of the hydraulic oil pressures in the cylinder chambers. A difference in pressure across the piston causes an imbalance of the forces on the piston, and the net force is transmitted to the bellcrank on the wing sleeve. For a given pressure differential, a slightly greater net force occurs when the actuator is extending than when it is retracting, due to the piston rod area. The force difference is small, however, (approximately five percent of the total force) and does not affect missile guidance.

By use of the bellcrank 94 FIG. 2, the translation of the piston is transmitted into rotary wing motion. The wing movement is limited to an incidence angle maximum of 20 (to each side of wing zero position), which corresponds to maximum piston travel of 2.12 inches for a full 40 wing movement.

As the piston moves in the actuator body, a corresponding movement of the actuating rod takes place in the feedback potentiometer 90. The potentiometer feeds back a voltage to the missile control system proportional to the amount of wing rotation from the wing zero position. The voltage is positive or negative depending upon the direction of wing rotation.

The overall operation of a servo valve and wing actuator combination is shown schematically in FIGS. 5a to 50.

A wing positioning signal, from the mixing section of the missile control system, is applied in the form of a differential current to the torque motor coils in the servo valve. The control arm is deflected in proportion to the current difference. If the movement of the control arm is to the left, the valve spool moves to the right, connecting high-pressure oil to the left actuator port and, at the same time, connecting the right actuator port to the low-pressure sump. The oil flow rate to and from the wing actuator is in proportion to the current differential applied to the torque motor coil.

The difierence in oil pressure across the actuator piston extends the piston, rotating the wing clockwise, as indicated by the arrow. When the wing moves from its zero position, the wiper arm 121 of the feedback potentiometer moves from its zero position and feeds a voltage back to the mixing section of the missile control system in a manner hereinafter described. The magnitude and sign of the feedback potentiometer voltage indicate wing position to the missile control system.

Each of the servo amplifiers, as hereinbefore stated, comprises a DC differential amplifier and a push-pull power output stage. The output of the push-pull stage is fed to the windings at 104 of the torque motor 911 which, in turn, activates the hydraulic system and causes the wings to move. The potentiometer is mechanically connected to each wing to furnish a feedback voltage to the servo amplifiers. This voltage acts as a position voltage. As the wings move a feedback voltage is picked off the potentiometers and fed through feedback networks to the grid of the servo amplifiers. When the desired wing positions are reached, the amplifiers are balanced, and there will be no further wing movement. The four servo amplifiers (one for each wing) are identical, therefore, the explanation will cover only the circuits for wing l.

A plus or minus feedback voltage from the servo wing potentiometer is applied to the signal grid of the dc. voltage amplifier tube through a suitable voltage divider network not shown, which is comprised of a series resistor and the grid circuit resistance (Rg) between the grid of the input amplifier tube and ground, which has a network gain of 0.0677. This feedback voltage is also applied to a voltage divider network comprised of two series connected resistors, which are connected between the plate of one tube of the power output stage and ground and has a network gain of 0.0415. The output of this network serves as an indication of wing position for telemetering. The plates of the push-pull connected output tubes and are connected to the +135 volt supply through a center tap of the field windings of the wing servo valves, which serves as a plate load for the tubes. The differential current of the two tubes will cause movement of the servo valve control arm or armature 118.

When an input voltage is applied to the grid of input tube of the direct coupled stage, the plate currents of the two triode tube sections of this stage becomes unbalanced. This causes a corresponding change in voltage applied to the respective grids of these tube stages. The differential current thus developed across the field windings of the servo valve causes a wing movement. As the wing moves in either direction from center position, the arm of the wing feedback potentiometer will be mechanically positioned so that when a positivegoing voltage is applied to the grid of a first tube section, a negative voltage will appear on the arm of the wing feedback potentiometer. When a negative voltage appears on the grid of this same triode tube section, the wing movement caused by the signal will cause the arm of the wing potentiometer to position itself so that a positive voltage appears on the arm of the wing potentiometer. When the desired wing position is reached, the feedback voltage is such that it balances the amplifier and there will be no further wing movement.

A boost phase follow-up control, WCDB, hereinafter described in greater detail is included in the system to maintain and stabilize the missile flight attitude within desired limits by sensing dispersion tendencies and by applying necessary control as the occasion or circumstances may necessitate to continuously maintain the missile within predetermined desired limits of the launching vector during the first stage of boost phase of missile flight.

So that subsequent details of the WCDB system and roll stabilization are clearly ascertainable, the energizing sequence of operation of various switches and relays in the electrical system before and after launching is now described. While the missile is situated on the loading machinery, the roll free gyro and WCDB two-axis free gyro are checked as being properly caged. If gyro caged indication is not present, a gyro caging command DC voltage is applied thereto from the external power. While on the launching machinery, the missile is warmed up from external power through a missile electrical contactor pad.

When the missile is positioned properly on the launcher and orientated for launching in the direction of attack external electrical power is fed through the launcher electrical contactor to the missile electrical contactor pad. The first step in energizing the missile is for the fire control officer to initiate a 26.5-volt D.C. command from external power through the electrical contactor pad which causes the roll free gyro and WCDB two-axis free gyro to uncage and causes through a set of closed contacts of a deactivated G- switch (not shown) a changeover from external power to missile battery power. The change over from external to internal power applies 26.5 VDC battery power to the normally open hydraulic pressure switch contact. The G-switch is a conventional thrust-operated, mechanical, self-latching device.

Upon completion of the uncaging of the gyros, the roll referenced position and the particular attitude of the missile as indicated by the orientation of the launcher is established. in the WCDB system, the particular attitude of the missile is determined by the twoaxis free gyro which places a referenced or prelaunch bias on capacitor 193 as explained in detail in the description of the attitude stabilization channel.

After power changeover has been made and with the completion of the uncaging sequence of the gyros, the 26.5 volt command is fed through a set of closed contacts of the deactivated G-switch energizing the hydraulic solenoid valve two seconds before booster ignition.

Energization of I the hydraulic solenoid valve activates to release hydraulic fluid into the missile hydraulic system which closes the hydraulic pressure switch. The closing of the hydraulic pressure switch causes the 26.5 volt battery power applied to the pressure switch contact through closed contacts of the deactivated G-switch to energize the activation relay 211 in the attitude stabilization channel of the WCDB system. The operation of the activation relay activates the WCDB system and closes the missile interlock to indicate that the missile is in ready-to-launch condition.

The missile ready indication to the missile fire control officer shows that the following operations within the missile are completed: roll free gyro uncaged, twoaxis free gyro uncaged, hydraulic system pressurized properly to provide power to the wing actuators, at-- titude stabilization channel activated to provide antideviation during boost phase. When these conditions exist, the launcher contactor is automatically removed and the booster ignited by electrical command through the booster forward launching shoes.

After launching, the initial acceleration of the missile causes the G"-switch to close and a mechanical lock,

not shown, keeps it closed. Upon activation of the G- switch, the following functions occur: (I) the power changeover line is interrupted, breaking the circuit to the missile electrical contactor pad; (2) battery power is supplied as a back up for the hydraulic solenoid valve, as latching voltage for the power changeover or transfer relay and to the activation relay to insure energized position during flight; (3) battery power is supplied to the inactivated separation relay shunting the hydraulic pressure switch, and (4) battery power is applied to the attitude stabilization channel to initiate the gain changer relay 212.

At separation of the booster from the missile, the separation microswitches 19 are allowed to activate which completes a ground circuit to the separation relay. The activation of the separation relay connects 26.5 VDC from' the battery power to the ramjet engine ignition unit and a separation command to the control system which activates relays to change contacts 70 from the WCDB system to the guidance control system 51.

The sensing apparatus for the wing control during boost system of the present invention includes three gyroscopes, 31, 32, 33 a rate conversion unit 34, and the attitude stabilization channel 35, as indicated in the block diagram of FIG. 3. The two axis free gyro 31 senses the amount of control necessary to minimize the dispersion of the missile in attitude during boost phase so that the missile has not deviated from the initial velocity vector of the missile at the end of the boost phase. The two-axis free gyro 31 provides two outputs proportional to the angular displacement about the inner and outer gimbals.

The A" and 8" plane rate gyros 32 and 33 indicate the rate of change in missile motion in each of the respective missile control planes which are determined by the location of the oppositely disposed pairs of missile wings. The output signals from the steering gyros 32 and 33 indicate only the rate of change'of missile angular orientation and are applied to the rate gyro conversion unit 34 in order to convert to desired D.C. signals.

The plane rate gyros 32 and 33 plus the various shaping networks or filters in the attitude stabilization channel 35 provide the necessary amplitude and phase characteristic as a function of frequency of the missile and its components so that the desired control of deviation of the missile can be achieved and so that necessary stability of the system and the missile can be maintained.

Referring now to FIG. 6, the steering rate gyro conversion unit 34 consists of a series circuit containing an amplifier stage, a cathode follower coupling state, and a bridge-type, germanium diode demodulator circuit 141.

The AC output of the rate gyro 33 is coupled to the control grid of amplifier 1318 through gain adjustment potentiometer 132 and coupling capacitor 133. Grid and cathode bias are provided by resistors 134, 135, and 136 to ground. Plate loading is provided by resistor 137 to the l35-volt electrical system supply. Cathode follower 131A, one half of the dual triode 131 is resistance-capacitance coupled by capacitor 138 from the plate of tube 131B and receives its control grid bias from resistor 139 connected to ground. Cathode follower 131A is used to isolate the impedances of the amplifier and demodulator stages.

The primary winding of transformer is connected to the cathode of cathode follower 131A (a 180 phase shift has occurred through amplifier 131B). Transformer 140 returns the phase of the signal to its original condition.

THE PRIMARY WINDING OF TRANSFORMER 3) IS CONNECTED TO THE CATHODE OF CATHODE FOLLOWER -0-A (A -*)PHASE SHIFT HAS OCCURRED THROUGH AMPLIFIER -0-B). TRANSFORMER RETURNS THE PHASE OF THE SIGNAL TO ITS ORIGINAL CONDITION.

The demodulator I41 employs silicon crystal diodes for rectification. Capacitors 161 and 163, and resistor 164 are components of a high-frequency filter network. The signal from the demodulator 141 is developed across load resistors 165 and 166 to ground.

The outputs of the rate gyro conversion unit are DC voltages varying in direct proportion to the change seen in the steering rate gyro due to missile motion or maneuver in the A and B planes. These outputs are applied to the attitude stabilization channel.

A schematic diagram of the attitude stabilization channel is shown in FIG. 7. This channel circuit is of two sections: (1) the three shaping networks isolated from each other by cathode followers, and (2) the timer-gain changes network. Two signal inputs are fed to each channel; only one, however, is shown in FIG. 7, because both channels are identical in operation. The two inputs are used signals from the two-axis (A and B planes) free gyro 31, and the respective signals from the A" or 8" plane rate gyros from the rate conversion unit 34.

The rate gyro conversion unit output is a steering rate gyro DC input to attitude stabilization channel. This steering rate gyro rectified signal enters attitude channel circuit through a low-pass RC filter comprised of resistors 172 and 173 and capacitor 174. This filter has highest pass characteristics at low frequencies. As frequency increases, the impedance of the filter to the frequency becomes greater. Its maximum attenuation occurs in the neighborhood of 550 c.p.s.

Cathode follower 175A is placed in the circuit between this low-pass filter and a following 42 c.p.s. notch filter, comprised of resistors 176 and 177 and capacitors I78 and 179. Cathode followers are used in this manner to isolate the impedances of filters, allowing the filters to operate independently of the signal.

Cathode follower 1758 is used to match the 42 c.p.s. notch filter and the following filter network. A dual notch filter composed of a parallel RC network, resistor 182 and capacitor 183, operates in series with a five-c component filter composed of resistors 184, 185, 186 and 187, and capacitor 188, to perform high-impedance filtration at the low frequency of approximately l to 3 c.p.s.

Tubes 189A and 1898 are two halves ofa dual triode tube 189. This tube and its associated circuitry make up a differential amplifier. The action of the differential amplifier will accept two DC signals of independently varying amplitude combine them through the conduction path of the tube, and will amplify the differential or sum of the two signals. The two signals presented to the differential amplifier are the amplified, demodulated,

and filtered DC signal from the steering rate gyro associated with the particular channel under discussion, and the DC signal from the two-axis free gyro 31. The modified steering rate gyro signal is presented on the grid of the leading triode 189A. The two-axis free gyro signal is presented on the grid of the second triode 189B. Differentiation of these two signals is achieved in the following manner: The signal from the steering rate gyro varies in accordance with missile motion along the mounting axis of the gyro. This signal causes a bias to appear on the grid of tube 189A, the leading triode. lf conduction should increase in tube 189A, a voltage drop appears across resistor 190 to ground. The voltage drop is reflected on the cathode of tube 189B, causing conduction to decrease to tube 189B. Decrease in conduction through tube 189B results in a drop across the plate load resistor 191, causing the voltage to increase on the plate of tube 1898. This signal is transferred to tube 192B cathode follower through capacitor 193 and a compensating resistor 194.

The 8" plane of the two-axis free gyro signal is brought into the circuit at the grid of 189B through isolation resistor 195. Potentiometer 186 and 196 are used to balance the two sections of the tube 189. Once balanced at a no signal level, any difference between the B plane rate gyro signal and the B plane free gyro signal will appear as the output to cathode follower 1928. This action is derived through degenerative feedback. Tube section 19213 is half of dual-triode tube 192. Section 192A is used similarly in channel A The output of cathode follower 192B appears across resistors 197 and 198 to ground. A portion of this signal is fed back to the grid of tube 18913 through resistor 199 and the potentiometer 196. Resistor 200 is used to increase the sensitivity of the potentiometer action, allowing more accurate adjustment of the balance of tube 189.

Before operation of the activation relay 211, with the missile on the launcher oriented for launching and with the two-axis free gyro in an uncaged status, a signal is applied to the grid of tube 18913 which places a referenced voltage signal on the plate of 1893. Since there exists short charging time constant path of low impedance through resistors 201 and 202 for capacitor 193, a charge is placed on capacitor 193 determined by the referenced voltage signal. The grid of cathode follower 192B is provided with biasing through potentiometer 201 through resistor 202 to the l35-volt source. This bias on the grid of 1928 is only applied before the operation of the activation relay. The bias is adjusted so that zero output of cathode follower 192B corresponds to the reference point for launch and that any variations of signals in either direction on the plate 1898 following launching of the missile pass through the capacitor 193 to the output of cathode follower 1923. Additionally, the bias provides isolation of any extraneous signals, which might appear before launching of the missile at the plate of 1898, from the output of 1928.

Upon operation of the activation relay 211, the circuit from grid of tube 1928 to resistors 201 and 202 is disconnected by open relay contacts of relay 211. After launching of the missile, the charge that was placed on blocking capacitor 193 before launching is maintained since there is no bleed-off resistor for the grid of 192B.

This condition establishes an exceedingly long time constant for the charge on capacitor 193 to bleed-off through tube 1923. Since the WCDB system operates for approximately 5 seconds, the ratio of the time constant of the channel circuit to the time of of operation of the system is about 200 to 1. With the ratio of time being so large, any slowly varying DC signals on the plate of tube 1898 after launching of the missile will pass through capacitor 193 to the output tube 192B of the WCDB system before there is any change in the charge on capacitor 193.

Two sets of contacts of relay 211 are used respectively for each channel. A third set of contacts is used as an interlock for the missile ready signal. The missile ready signal will not be passed to the external power connector until the activation relay is energized.

After 1.5 seconds of boost flight, the increased speed of the missile-booster combination enhances the effectiveness of the wing surfaces and therefore a smaller control signal from the WCDB system to the mixer amplifier 71 is required to hold the missile-booster combination on a path parallel to the launching aiming vector for remainder of the boost phase. The gain of the channel is reduced to approximately one-third its maximum value after 1.5 seconds of boost flight for the remainder of boost phase of missile flight, that is, until missile-booster separation occurs. Gain changing is done by removing resistor 197 from the series circuit consisting of resistors 197 and 198 in series. Resistor 197 is removed by a contact of gain changer relay 213. Tube 221 is a gas-filled tetrode employed in a 1.5 second delay circuit. The timer activation signal to energize relay 212 is a 26.5 VDC from the battery, routed through the G-switch to relay 212 when the G"-switch is activated at launch. Before activation of relay 212, the grid of timer tube 221 carries 1 35 volts, cutting off any possibility of conduction. At launching of the missile, this -1 35 volts is removed from the grid through the action of a contact of relay 212. At the same time. another contact of relay 212 applies volts to the grid of timer tube 221. The tube now would tend to conduct heavily; however, a relatively large valued capacitor 222 is placed in the grid cathode circuit. The time delay in discharging and recharging this condenser 222, plus the inherent delay of the gas-filled tube 221, causes an overall delay of 1.5 seconds before maximum conduction occurs in the tube 221. The coil of relay 213 is connected in series with the plate of the timer tube 221. The +135 volt potential appears on the coil of relay 213 when relay 212 is energized; however, no current can flow until the gas tube 221 begins conduction. At conduction, gain changer relay 213 is energized and is used for switching the output gain in both A and B channels of the attitude stabilization channel 35. At 1.5 seconds after launching of the missile, two-thirds of a series load resistance to ground in the output tube 192 is shorted out in both channels A and B." The WCDB system output signals from the attitude stabilization channel 35 in either the A or B plane are combined with the roll stabilization channel signals in the mixer amplifiers 71.

Since the missile steers itself on the assumption that missile orientation is maintained with the planes of the wings making 45 angles with a vertical plane through the missile axis, a roll stabilization system is necessary to correct for any deviation from this roll referenced position.

By a foregoing description of the sequence of operation of the various relays while the vehicle is on the launcher, the roll stabilization system is activated immediately before launching so that the missile is in the upside up position after launching to enable capture by the narrow guidance radar beam at the earliest. possible time after the boost phase of missile flight.

In the block diagram of the roll stabilization system, of FIG. 8, an output signal from the roll free gyro 61 which measures the amount and direction of missile roll from the reference position is limited electrically by the limiter circuit 62 to a nominal value ofi20. The roll rate gyro 63 output signal indicates the rate and direction of roll of the missile. The function of the limiter 62 is to allow an increase of the ratio of rate-toposition signal for better performance during roll maneuvers. The roll error and the roll rate error signals from the roll free gyro and the roll rate gyro are resistively mixed and the combined signal amplified in the roll stabilization channel 64 by amplifier 65. The gain of amplifier 65 is varied by static pressure switches as a function of ambient pressure. This is accomplished by a ram-static pressure probe 54 sensing the static pressure to operate the static pressure switches 53 to change the gain of amplifier 65 at various altitudes to compensate for wing effectiveness with altitude. The output of amplifier 65 is connected through cathode follower 66 to the demodulator 67. The demodulator converts the combined signals from the roll free and roll rate gyros to a DC voltage which is connected through an output cathode follower 68 to the mixer amplifiers 71. The output signals from the roll stabilization channel result in differential movement of the wings in the same plane to limit missile roll.

The incorporation of the WCDB system in the missile control system to sense and indicate signals to control the attitude and direction of the missile relative to the launching vector in combination with a roll sta' bilization system during boost phase of missile flight establishes missile flight stability or performance for an aerodynamically unstable missile and additionally overcomes the dispersion influences caused by cross winds or fin misalignments enabling capture of the missile in an external narrow wave of electromagnetic energy.

The present invention could be used in other types of guidable vehicles or missiles such as a missile propelled by a dual thrust single chamber rocket or a two stage rocket. Additionally, the present system could be used in a homing type of guidable missile wherein the sensing device for indicating the pitch and yaw of the missile and rate of change thereof would measure the deviation from a predetermined target homing reference once the missile is in terminal or homing phase of flight. The sensing apparatus of the boost control system is well adaptable as a damping or limiting means for the pitch and yaw of a missile during the guidance phase of flight.

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.

What is claimed is:

l. A control system for steering a zero velocity launched guidable tandem coupled booster missile airborne vehicle during the boost phase of flight; comprising means for sensing the deviation of the vehicle in flight to provide steering signals, means for actuating the vehicle aerodynamic surfaces, means for supplying power to said actuating means, said power supply means being precharged before launching of the vehicle whereby power is available immediately upon the launching of the vehicle to operate said actuating means during the beginning portion of the boost phase of flight, said power being coupled to said actuating means, said sensing means comprising means for measuring the magnitude and direction of vehicle roll for providing roll stabilization steering signals, means for determining the attitude of the vehicle from the launching vector during boost phase and establishing boost phase steering signals, means for mixing the roll stabilization steering signals with the boost phase steering signals, said steering signals from said mixing means being applied to said actuating means to operate the vehicle aerodynamic surfaces thereby changing the attitude of the vehicle and the direction of vehicle flight by movement of said vehicle aerodynamic surfaces, whereby vehicle roll stabilization and vehicle attitude stability is provided throughout the boost phase of flight enabling capture of the missile in a narrow radar beam at the end of boost phase of flight.

2. A missile control system for wing-steering ofa missile-booster combination during the boost phase and the ram-jet missile during guidance phase of missile flight; comprising means for sensing the deviation of the missile in flight to provide steering signals, means for actuating the movement of the missile wings, said actuating means comprising a wing servo system having a servo valve, a wing actuator and a feedback potentiometer for each missile wing, said feedback potentiometer providing a feedback signal being applied to said sensing means for representing the missile wing positions, means for supplying hydraulic power to said wing servo system, said hydraulic power being coupled to said servo valve and being connected through said servo valve to said wing actuator, said sensing means comprising means measuring the missile roll from a pre-launch roll reference position for providing roll stabilization steering signals, means for determining the attitude of the missile-booster combination from the launching vector during boost phase and establishing boost phase steering signals, means receiving intelligence during the guidance phase for providing guidance phase steering signals, means for switching from the boost to guidance steering signals, means for mixing said roll stabilization steering signals with said phase steering signals from said switching means, said steering signals from said mixing means being applied to said servo valve to regulate the hydraulic power used to operate said wing actuator thereby changing the position of the missile wings causing an attitude deviation correction of the missile, and a change of missile flight direction whereby the missile is roll stabilized throughout missile flight, missile-booster stability is provided during boost phase of missile flight thereby overcoming adverse influences on the missile from the launching attitude of the missile and intelligence control is provided during guidance phase of missile flight.

Claims (13)

1. A control system for steering a zero velocity launched guidable tandem coupled booster missile airborne vehicle during the boost phase of flight; comprising means for sensing the deviation of the vehicle in flight to provide steering signals, means for actuating the vehicle aerodynamiC surfaces, means for supplying power to said actuating means, said power supply means being precharged before launching of the vehicle whereby power is available immediately upon the launching of the vehicle to operate said actuating means during the beginning portion of the boost phase of flight, said power being coupled to said actuating means, said sensing means comprising means for measuring the magnitude and direction of vehicle roll for providing roll stabilization steering signals, means for determining the attitude of the vehicle from the launching vector during boost phase and establishing boost phase steering signals, means for mixing the roll stabilization steering signals with the boost phase steering signals, said steering signals from said mixing means being applied to said actuating means to operate the vehicle aerodynamic surfaces thereby changing the attitude of the vehicle and the direction of vehicle flight by movement of said vehicle aerodynamic surfaces, whereby vehicle roll stabilization and vehicle attitude stability is provided throughout the boost phase of flight enabling capture of the missile in a narrow radar beam at the end of boost phase of flight.

2. A missile control system for wing-steering of a missile-booster combination during the boost phase and the ram-jet missile during guidance phase of missile flight; comprising means for sensing the deviation of the missile in flight to provide steering signals, means for actuating the movement of the missile wings, said actuating means comprising a wing servo system having a servo valve, a wing actuator and a feedback potentiometer for each missile wing, said feedback potentiometer providing a feedback signal being applied to said sensing means for representing the missile wing positions, means for supplying hydraulic power to said wing servo system, said hydraulic power being coupled to said servo valve and being connected through said servo valve to said wing actuator, said sensing means comprising means measuring the missile roll from a pre-launch roll reference position for providing roll stabilization steering signals, means for determining the attitude of the missile-booster combination from the launching vector during boost phase and establishing boost phase steering signals, means receiving intelligence during the guidance phase for providing guidance phase steering signals, means for switching from the boost to guidance steering signals, means for mixing said roll stabilization steering signals with said phase steering signals from said switching means, said steering signals from said mixing means being applied to said servo valve to regulate the hydraulic power used to operate said wing actuator thereby changing the position of the missile wings causing an attitude deviation correction of the missile, and a change of missile flight direction whereby the missile is roll stabilized throughout missile flight, missile-booster stability is provided during boost phase of missile flight thereby overcoming adverse influences on the missile from the launching attitude of the missile and intelligence control is provided during guidance phase of missile flight.

3. In a missile control system for steering a missile-booster combination during the boost phase and a missile during the guidance phase of missile flight by movement of the missile wings; comprising missile wings, means for sensing the deviation of the missile in flight to provide electronic steering signals for regulating the movement of the missile wings, means for actuating the movement of the missile wings, feedback means connected to the missile wings and to said sensing means for providing a feedback signal for said sensing means which is indicative of missile flight position, said sensing means comprising means for measuring the amount and direction of missile roll from a pre-launch roll reference position providing electronic roll stabilization steering signals, means for determining the attitude of the missile-booster combination froM the launching vector during boost phase establishing electronic boost steering signals, means for receiving intelligence during the guidance phase providing electronic guidance steering signals, means for switching from the electronic boost steering signals to the guidance steering signals, means for mixing and amplifying the electronic roll stabilization steering signals with the steering signals from said switching means, said feedback signal from said feedback means being connected to said mixing and amplifying means, said electronic steering signals from said mixing and amplifying means being applied to said actuating means to regulate the movement of the missile wings by said actuating means thereby changing the attitude of the missile and the direction of missile flight whereby roll stabilization is provided throughout missile flight, missile-booster stability is provided during boost phase and intelligence control is provided during guidance phase of missile flight.

4. A control mechanism for steering a missile-booster combination during boost phase of missile flight by movement of the missile wings; comprising missile wings, means for sensing the deviation of the missile in flight to provide electronic steering signals for changing the flight position of the missile by regulating the movement of the missile wings, means for actuating the missile wings, said actuating means comprising a wing servo system having a servo valve, a wing actuator and a feedback potentiometer for each wing, means for supplying hydraulic power to said actuating means, said hydraulic power supply being coupled to said servo valve and being connected through said servo valve to said wing actuator, said sensing means comprising means measuring the missile roll for providing electronic roll stabilization steering signals, means determining the attitude deviation of the missile-booster combination relative to the launching vector during boost phase for establishing electronic boost phase steering signals, means for mixing and amplifying said electronic roll stabilization steering signals with said boost phase steering signals, said feedback potentiometer being electrically connected to said mixing and amplifying means, and means for mechanically connecting said feedback potentiometer to said missile wings, said feedback potentiometer being operated by said connecting means for developing a signal voltage indicative of the missile wing positions and providing a feedback signal to said mixing and amplifying means, said electronic steering signals being applied from said mixing and amplifying means to said servo valve to operate said wing actuator thereby changing the position of the missile wings causing a change of attitude of the missile and the direction of missile flight whereby the missile is roll stabilized throughout boost phase of flight, and missile-booster stability is provided during boost phase of missile flight.

5. The control mechanism as defined in claim 4 wherein said hydraulic power supply means includes means precharged to a predetermined pressure before launching of the missile-booster combination whereby hydraulic power is available immediately upon the launching of the missile-booster combination to operate said wing actuator during the beginning portion of the boost phase of missile flight.

6. The mechanism as defined in claim 5 wherein said hydraulic power supply means includes means responsive to diffuser air pressure during boost phase for supplementing and/or supplanting pressurization from said precharged pressure means.

7. The control mechanism as defined in claim 6 wherein said sensing means comprises means for receiving intelligence after boost phase of missile flight and for determining the missile deviation from the radar guidance beam during guidance phase for establishing electronic guidance phase steering signals, means for switching from said electronic boost phase steering signals to said guidance phase steering signals after boost phase of missile flight, and said mixing and amplifying means combining said electronic guidance phase from said steering signals switching means and said electronic roll stabilization steering signals after boost phase of flight whereby external intelligence control is provided after boost phase of missile flight.

8. In a system of the character described for control of the boost phase of missile flight, the combination; a guidable tandem coupled booster missile, said tandem booster being separable from said missile after boost phase of missile flight, said booster and said missile having a plurality of fixed aerodynamic surfaces, said missile with said tandem coupled booster being aerodynamically unstable, movable missile wing surfaces, means for providing actuation and control of said wing surfaces during the boost phase of missile booster flight, means for sensing flight deviation from a predetermined initial flight pattern and developing correctional intelligence signals correlative thereto, means for determining missile roll deviation from a pre-launch roll reference position and developing correctional roll stabilization signals correlative thereto and means for effecting stabilization control of said missile applying said correctional intelligence and roll stabilization signals to said wing actuating means.

9. In a system for control of a missile-booster vehicle during boost phase of flight, the combination of a missile-booster vehicle comprising a guidable missile and a separable tandem coupled booster, said booster and said missile having fixed aft positioned fins respectively, said missile having movable wings disposed at the longitudinal center of gravity of the missile, said missile-booster vehicle being aerodynamically unstable, and a boost control system within said missile, said system having means for sensing and establishing electronic steering signals indicative of the pitch and yaw and rate of change thereof of said missile-booster vehicle relative to the launching vector thereof, wing actuating means and means applying said steering signals to said wing actuating means to move said missile wings to insure aerodynamically stable flight and attitude stability of said missile-booster vehicle during boost phase of flight until said booster separates from the missile.

10. A control mechanism of a zero velocity launched guidable missile with tandem coupled booster for controlling the flight of the missile-booster vehicle during the boost phase of vehicle flight; comprising a boost control system for sensing and establishing electronic boost phases steering signals indicative of the pitch and yaw and rate of change thereof of the missile-booster vehicle relative to the launching vector thereof, a wing servo system for actuating the movement of the missile wings in response to said boost phase steering signals, means within said boost control system for programming the gain of said boost phase steering signals during boost phase acceleration of said vehicle to insure effective control of the movement of missile wings whereby attitude stability of the missile-booster vehicle is continuously maintained during boost phase of vehicle flight and capture of the missile in a narrow guidance radar beam is made possible at the end of boost phase of vehicle flight upon separation of the booster from the missile.

11. The control mechanism as recited in claim 1 wherein said mechanism includes a roll stabilization system for determining and establishing roll stabilization signals indicative of the roll and rate of change thereof of the missile-booster vehicle relative to a pre-launch roll reference position, said wing servo system being actuated in response to said roll stabilization signals for movement of the missile wings whereby roll stabilization of the missile-booster vehicle is continuously maintained during boost phase of vehicle flight.

12. In a control mechanism for enhancing flight performance of a missile-booster combination of an aerodynamically unstable configuration bY movement of the wings of a ram-jet missile with a tandem booster during boost phase of missile flight; comprising a boost control system for determining the attitude of the missile-booster combination relative to the launching vector of the missile-booster combination establishing electronic boost phase steering signals, amplifying means for said steering boost phase signals and roll stabilization circuits for measuring the missile-booster combination roll providing electronic roll stabilization steering signals, said amplifying means combining said boost phase steering signals with said roll stabilization signals whereby said steering signals regulate the movement of missile wings providing missile-booster combination stability and roll stabilization during boost phase of missile flight enabling capture by an external wave of electromagnetic energy after separation of the missile-booster combination.

13. A missile control system for steering an aerodynamically unstable missile-booster combination during the boost phase of missile flight by movement of the missile wings; comprising means for sensing the deviation of the missile in flight to provide electronic steering signals for changing the position of the missile in flight by regulating the movement of the missile wings, means for actuating the movement of the missile wings, said actuating means comprising a wing servo system for each wing having a servo valve, a wing actuator and a feedback potentiometer, means for supplying hydraulic power to said actuating means, said hydraulic power being coupled to said servo valve and being connected through said servo valve to said wing actuator, said hydraulic power supply means being precharged before launching of the missile-booster combination whereby hydraulic power is available immediately upon launching of the missile-booster combination to operate said wing actuators during at least the beginning portion of boost phase of missile flight, said sensing means comprising means for measuring the missile roll providing electronic roll stabilization steering signals, means for determining the attitude of the missile-booster combination from the launching vector during boost phase establishing electronic boost steering signals, means for mixing and amplifying said electronic roll stabilization steering signals with said boost steering signals, said feedback potentiometer being operatively connected to said mixing and amplifying means, and means for driving said feedback potentiometer from said missile wing actuator, said feedback potentiometer providing a feedback signal to said mixing and amplifying means indicative of the instantaneous missile wing position, said electronic steering signals being applied from said mixing and amplifying means to said servo valve to regulate the amount of hydraulic power used to operate said wing actuator thereby changing the position of the missile wings causing a change of attitude of the missile and the direction of missile flight whereby the missile booster combination is roll stabilized through boost phase of missile flight, and missile-booster stability is provided during boost phase of missile flight, thereby overcoming deviations of the missile-booster combination from the launching vector due to adverse influences resulting from cross winds or misalignments of the aerodynamic surfaces of the missile permitting capture of the missile in a narrow guidance radar beam after boost phase of missile in a narrow guidance radar beam after boost phase of missile flight upon separation of the booster from the missile.

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