US3094072A - Aircraft, missiles, missile weapons systems, and space ships - Google Patents

Aircraft, missiles, missile weapons systems, and space ships Download PDF

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US3094072A
US3094072A US701571A US70157157A US3094072A US 3094072 A US3094072 A US 3094072A US 701571 A US701571 A US 701571A US 70157157 A US70157157 A US 70157157A US 3094072 A US3094072 A US 3094072A
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cowl
thrust
nozzle
pressure
plug
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Arthur R Parilla
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    • FMECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F42AMMUNITION; BLASTING
    • F42BEXPLOSIVE CHARGES, e.g. FOR BLASTING, FIREWORKS, AMMUNITION
    • F42B10/00Means for influencing, e.g. improving, the aerodynamic properties of projectiles or missiles; Arrangements on projectiles or missiles for stabilising, steering, range-reducing, range-increasing or fall-retarding
    • F42B10/60Steering arrangements
    • F42B10/66Steering by varying intensity or direction of thrust
    • F42B10/665Steering by varying intensity or direction of thrust characterised by using a nozzle provided with at least a deflector mounted within the nozzle
    • BPERFORMING OPERATIONS; TRANSPORTING
    • B64AIRCRAFT; AVIATION; COSMONAUTICS
    • B64CAEROPLANES; HELICOPTERS
    • B64C9/00Adjustable control surfaces or members, e.g. rudders
    • B64C9/38Jet flaps
    • BPERFORMING OPERATIONS; TRANSPORTING
    • B64AIRCRAFT; AVIATION; COSMONAUTICS
    • B64GCOSMONAUTICS; VEHICLES OR EQUIPMENT THEREFOR
    • B64G1/00Cosmonautic vehicles
    • B64G1/22Parts of, or equipment specially adapted for fitting in or to, cosmonautic vehicles
    • B64G1/40Arrangements or adaptations of propulsion systems
    • B64G1/401Liquid propellant rocket engines
    • BPERFORMING OPERATIONS; TRANSPORTING
    • B64AIRCRAFT; AVIATION; COSMONAUTICS
    • B64GCOSMONAUTICS; VEHICLES OR EQUIPMENT THEREFOR
    • B64G1/00Cosmonautic vehicles
    • B64G1/22Parts of, or equipment specially adapted for fitting in or to, cosmonautic vehicles
    • B64G1/40Arrangements or adaptations of propulsion systems
    • B64G1/403Solid propellant rocket engines
    • FMECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F02COMBUSTION ENGINES; HOT-GAS OR COMBUSTION-PRODUCT ENGINE PLANTS
    • F02CGAS-TURBINE PLANTS; AIR INTAKES FOR JET-PROPULSION PLANTS; CONTROLLING FUEL SUPPLY IN AIR-BREATHING JET-PROPULSION PLANTS
    • F02C7/00Features, components parts, details or accessories, not provided for in, or of interest apart form groups F02C1/00 - F02C6/00; Air intakes for jet-propulsion plants
    • F02C7/04Air intakes for gas-turbine plants or jet-propulsion plants
    • F02C7/042Air intakes for gas-turbine plants or jet-propulsion plants having variable geometry
    • FMECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F02COMBUSTION ENGINES; HOT-GAS OR COMBUSTION-PRODUCT ENGINE PLANTS
    • F02KJET-PROPULSION PLANTS
    • F02K9/00Rocket- engine plants, i.e. plants carrying both fuel and oxidant therefor; Control thereof
    • F02K9/08Rocket- engine plants, i.e. plants carrying both fuel and oxidant therefor; Control thereof using solid propellants
    • F02K9/32Constructional parts; Details
    • F02K9/38Safety devices, e.g. to prevent accidental ignition
    • FMECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F02COMBUSTION ENGINES; HOT-GAS OR COMBUSTION-PRODUCT ENGINE PLANTS
    • F02KJET-PROPULSION PLANTS
    • F02K9/00Rocket- engine plants, i.e. plants carrying both fuel and oxidant therefor; Control thereof
    • F02K9/80Rocket- engine plants, i.e. plants carrying both fuel and oxidant therefor; Control thereof characterised by thrust or thrust vector control
    • FMECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F02COMBUSTION ENGINES; HOT-GAS OR COMBUSTION-PRODUCT ENGINE PLANTS
    • F02KJET-PROPULSION PLANTS
    • F02K9/00Rocket- engine plants, i.e. plants carrying both fuel and oxidant therefor; Control thereof
    • F02K9/80Rocket- engine plants, i.e. plants carrying both fuel and oxidant therefor; Control thereof characterised by thrust or thrust vector control
    • F02K9/86Rocket- engine plants, i.e. plants carrying both fuel and oxidant therefor; Control thereof characterised by thrust or thrust vector control using nozzle throats of adjustable cross- section
    • FMECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F02COMBUSTION ENGINES; HOT-GAS OR COMBUSTION-PRODUCT ENGINE PLANTS
    • F02KJET-PROPULSION PLANTS
    • F02K9/00Rocket- engine plants, i.e. plants carrying both fuel and oxidant therefor; Control thereof
    • F02K9/80Rocket- engine plants, i.e. plants carrying both fuel and oxidant therefor; Control thereof characterised by thrust or thrust vector control
    • F02K9/92Rocket- engine plants, i.e. plants carrying both fuel and oxidant therefor; Control thereof characterised by thrust or thrust vector control incorporating means for reversing or terminating thrust
    • FMECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F05INDEXING SCHEMES RELATING TO ENGINES OR PUMPS IN VARIOUS SUBCLASSES OF CLASSES F01-F04
    • F05DINDEXING SCHEME FOR ASPECTS RELATING TO NON-POSITIVE-DISPLACEMENT MACHINES OR ENGINES, GAS-TURBINES OR JET-PROPULSION PLANTS
    • F05D2240/00Components
    • F05D2240/10Stators
    • F05D2240/12Fluid guiding means, e.g. vanes
    • F05D2240/128Nozzles
    • F05D2240/1281Plug nozzles
    • YGENERAL TAGGING OF NEW TECHNOLOGICAL DEVELOPMENTS; GENERAL TAGGING OF CROSS-SECTIONAL TECHNOLOGIES SPANNING OVER SEVERAL SECTIONS OF THE IPC; TECHNICAL SUBJECTS COVERED BY FORMER USPC CROSS-REFERENCE ART COLLECTIONS [XRACs] AND DIGESTS
    • Y02TECHNOLOGIES OR APPLICATIONS FOR MITIGATION OR ADAPTATION AGAINST CLIMATE CHANGE
    • Y02TCLIMATE CHANGE MITIGATION TECHNOLOGIES RELATED TO TRANSPORTATION
    • Y02T50/00Aeronautics or air transport
    • Y02T50/40Weight reduction
    • Y02T50/42Airframe
    • Y02T50/44Design measures

Description

June 18, 1963 A. R. PARILLA AIRCRAFT, MISSILES, MISSILE WEAPONS SYSTEMS, AND SPACE SHIPS 5 Sheets-Sheet 1 Filed Dec. 9, 1957 FIG. IA

INVENTOR. ARTHUR R. PARILLA BY ww ATTORNEYS llllll'!" June 18, 1963 A. R. PARILLA 3,094,072

AIRCRAFT, MISSILES, MISSILE WEAPONS SYSTEMS, AND SPACE SHIPS Filed Dec. 9, 1957 5 Sheets-Sheet 2 57 55 5655 Amy u I I F l G- 9 F I G- 8 2&3 52 2/2 203 205 52 25 20 85 206 fl 205 8B INVENTOR ARTHUR R. PARILLA 0; y 4 z f tq 40"; L /%//M,Fmyam,- a/Z m m ATTCD RNEYS June 18, 1963 Filed Dec. 9, 1957 R. PARILLA AIRCRAFT, MISSILES, MISSILE WEAPONS SYSTEMS, AND SPACE SHIPS 5 Sheets-Sheet 5 INVENTOR. RV PARILLA ARTHLJ R ATTCDRNEYS June 18, 1963 A. R. PARILLA 3,094,072

AIRCRAFT, MISSILES, MISSILE WEAPONS SYSTEMS, AND SPACE SHIPS Filed Dec. 9, 1957 5 Sheets-Sheet 4 F I 3. I45

M M2 I l G- |4C Z5 THRUST 50057512 CASE ARM /Z*-#7% THRUST TERM/NATION SEPARATION T DESTRUCTION PRIMARY THROAT TIME-DELAY Eco/V05 AREA SW/TCH wens/1.95s

THROAT AREA DECREASES SECONDARY SECONDARY F I TIME-DELAY /a/v/7/0/v SWITCH m5 7/ i 779 r I 175 /74 m 776 /77 i 5 w F I 3. I?

INVENTOR.

u ARTHUR R. RARILLA 5 BY E R TIME t, f; ATTO R N EYS June 18, 1963 PARILLA SYSTEMS, AND SPACE SHIPS 5 Sheets-Sheet 5 Filed Dec. 9, 1957 7 0 0 0 0 3 3 4 0 0/ 9 1 0 W lie) 0 F n \II\\LI\\ 0 0 M 2 Wm o y 3 W J [J M W Z 0/ 3 0/ 00 0 5 0 i 0 3 3 l. 3 0, 4 j y FIG- INVENTOR. ARTHUR R. PAR! l A flay, Fuuym, u/ZM 4 AT TO R N E vs United States Patent $394,072 AIRCRAFT, MISSILES, MllSlLE WEAPONS SYSTEMS, AND SPACE SEES Arthur R. Pariila, 34 Crestview Road, Mountain Lakes, NJ. Filed Dec. 9, 1957, Ser. No. 761,571 27 Claims. (Ql. l02--Sll) This invention relates to improvements in aircraft, missiles, missile weapon systems and space ships. It is more specifically related to improvements in propulsion systems of both the rocket and air-breathing type, which will extend the operational limits of such vehicles; and provide greater accuracy and range through improved flexibility and control over engine operation together with improved engine efliciency.

It is the purpose of this invention to provide simple means for varying the throat area of supersonic nozzles for jet propulsion engines; to provide variable area expansion ratios for such nozzles; to incorporate within the nozzle itself simple means for thrust vector directional control, and thrust termination; and to integrate these features with the variable area geometry; to adapt these principles to variable geometry inlets for air-breathing engines; and to achieve maximum performance at all altitudes including space flight.

It is a further purpose of this invention to integrate these new concepts on propulsion systems with the operational use of new missiles; to provide new concepts in anti-space missile weapons systems which will provide greater accuracy and reliability for defense against attack by objects travelling through space; and to provide improved anti-ballistic missile weapons systems for local defense.

When applied to solid propellant rocket engines, the new principles make possible a controllable solid propellant rocket engine whose thrust may be varied in both magnitude and direction at will during flight, including controlled thrust termination; and make possible substantial reduction in weight of inert parts beyond that obtainable by use of higher strength materials. Thus, it adds to the proven reliability of the solid propellant rocket engine a new flexibility of engine operation which equals and surpasses that of present conventional liquid propellant rocket engines.

When applied to liquid propellant rocket engines, improved thrust vector-directional control may be provided in which the thrust chamber is rigidly mounted to the missile, simplifying plumbing for greater reliability, and eliminating the weight of heavy gimbal mounted thrust structures. It improves throttleability over a broader thrust range.

When applied to air-breathing engines, either turbo-jet, ram-jet, or ducted rockets, variable geometry inlets and exits permit operation over a wider range of flight mach numbers, angle of attack and yaw; improved maneuverability, rate of climb, acceleration, and fuel economy. Thrust vector control, as well as variable nozzles areas, may also be used for improved maneuverability and stability of aircraft while reducing the size of aerodynamic control surfaces, or eliminating them.

Space ships, satellites, and upper stages of long range ballistic missiles which are propelled beyond the Earths atmosphere may be designed to operate most efficiently in vacuum conditions. As the pressure ratios approach infinity, nozzles with extremely large area expansion ratios may be provided for maximum propellant specific impulse, without penalizing nozzle weight or overall length as with conventional nozzles. These advantages may be gained while retaining thrust vector control, variable thrust, and other features.

It is, therefore, the purpose of this invention to advance the state of the art by accomplishing the following objects:

(1) A primary object of this invention is to apply new principles of fluid flow to the design of jet engine nozzles, adaptable to provide a variable throat area.

(2) It is a further object to provide a nozzle for jet engines whose expansion ratio may be varied during flight.

3) Another object is to provide a nozzle for jet engines incorporating novel means for thrust vector directional control.

(4) Another object is to provide a nozzle for jet engines incorporating novel means for controlled thrust termination upon command under any flight condition.

(5) Another object is to provide improved nozzles for jet engines having very large area expansion ratios for operation at high altitudes and in space, while permitting a reduction in nozzle weight and length.

(6) Another object is to provide a nozzle for jet engines incorporating novel means whereby variable throat area, variable expansion ratio, thrust vector directional control, and thrust termination may be integrated within the same nozzle, or in various combinations.

(7) Another object is to provide a nozzle for jet engines in which the variable throat area operates automatically responsive to variation in mass flow rate of the working fluid.

(8) Another object is to reduce the weight of jet engines by reducing the weight of the nozzle, thrust vector control system, means for thrust termination, and mounting structure.

(9) Another object is to improve the reliability of jet engines by reducing the number of subsidiary components and incorporating their same functions within the nozzle itself.

(10) Another object is to provide a controllable solid propellant rocket engine which operates at substantially constant pressure independent of the temperature sensitivity of the propellant; progressivity and/ or regressivity of the burning surface of the propellant grain; and/ or erosive burning.

(11) Another object is to reduce the Weight of solid ropellant rocket cases beyond the use of higher strength materials, by providing substantially constant pressure operation, thereby reducing or eliminating peak pressure design requirements.

(12) Another object is to provide a solid propellant rocket engine in which the variation in thrust due to ambient temperature changes or variable burning surfaces are minimized.

(13) Another object is to provide a controllable solid propellant rocket engine in which the magnitude of thrust may be varied during flight in any manner as desired.

(14) Another object is to provide for a disposable booster rocket case which may be destroyed in air after separation of missile from booster, and before impact on friendly territory.

(15) Another object is to provide dual stage solid propellant rocket engines capable of operating at extreme ratics of maximum to minimum thrust levels, for boost (acceleration) and sustained (constant velocity) propulsion of missiles.

(17) Another object is to provide mechanical or fluid springs for automatically controlling the variable throat area of jet engine nozzles.

(18) Another object is to provide fluid, and/or electrically operated control systems and actuators for controlling the variable throat area of jet engine nozzles.

(19) Another object is to provide improved ballistic missiles whose flight trajectory is related to engine performance so as to provide optimum nozzle thrust co eflicient as function of altitude by automatically varying the nozzle area expansion ratio.

(20). Another object is to broaden the operating range of turbo-jet engines, ram-jet engines, ducted rocket engines, or other air-breathing engines by providing an improved nozzle with variable throat area, and variable expansion ratios.

(21) Another object is to broaden the operating range of air breathing engines such as turbo-jets, ram-jets, ducted rockets, over wider range of flight mach numbers and flight attitudes, such as angles of attack and yaw, by providing variable geometry inlets.

(22) Another object is to improve the maneuverability and controllability of aircraft and missiles powered with air breathing engines, sucuh as turbo-jet, ram-jet, ducted rockets and the like, by utilizing novel means for jet engine thrust vector directional control, thereby reducing the size ,or eliminating, aerodynamic control surfaces.

(23) Another object is to improve the reliability of ram-jet engines by utilizing the proven reliability of solid propellant fuels together with the controllable features of variable geometry inlets and exits.

(24) Another object is to provide novel configurations for rocket engine nozzles operating in the vacuum conditions of space flight.

These and other objects will become apparent from the following detailed description, read in connection with the annexed drawings, in which similar reference characters represent similar parts, and in which:

IGURES 1A, B and C are fragmentary views in longitudinal section of an axi-sy-mmetric nozzle employing external supersonic expansion based on the principle of Prandtl-Meyer flow around a corner; FIGURE 1A illustrates a plug nozzle having an isentropic surface; FIGURE 1B, a single conical plug; and FIGURE 1C, a multiple conical plug.

FIGURE 2 is a fragmentary view in longitudinal section of a variable throat area plug nozzle in accordance with this invention showing the same nozzle principle in FIGURE 1A but adapted to provide a variable throat area with the balanced pressure forces acting aft, that is, to the right in FIGURE 2.

FIGURE 3 is a fragmentary view in longitudinal section of a variable throat area plug nozzle in accordance with this invention showing the same principle adapted to pro vide a variable throat are, but with the unbalanced pressure forces acting forward, that is, to the left as viewed in FIGURE 3.

FIGURE 4 is a fragmentary view in longitudinal section illustrating a nozzle incorporating thrust vector directional control.

FIGURE 5A is a fragmentary view in longitudinal section of an embodiment of nozzle in accordance with this invention having provision for thrust termination.

FIGURE 5B is a fragmentary view in longitudinal section of a nozzle embodiment of this invention showing alternate means for thrust termination.

FIGURE 6 is a fragmentary view in longitudinal section of a nozzle embodiment of this invention incorporating means for providing combined variable area, thrust vec tor control, and thrust termination.

- FIGURE 7 is a fragmentary view in longitudinal section showing an alternate form of seal for the cowl of FIG- URE 1x1. i

FIGURE 8 is a longitudinal view of a modified embodiment of a plug nozzle in accordance with this invention incorporating means for providing combined variable area and thrust vector control without a gimbal ring.

FIGURE 8B is an elevational view taken on the line 8B8 B of FIGURE 8 showing a detail of the plug nozzle of FIGURE 8.

FIGURE 9 is a fragmentary view in longitudinal section of a modified embodiment of a nozzle showing means for providing variable area, thrust vector control and thrust termination with nozzle employing only internal expansion.

FIGURE 10 is a view in longitudinal section of a solid 4 propellant rocket engine embodying a variable throat area plug nozzle with thrust vector control and thrust termination in accordance with this invention.

FIGURE 11 is a diagram illustrating effect of variable throat area nozzle on the internal ballistics of solid propellant rocket engines.

FIGURE 12 is a fragmentary view in longitudinal section showing as another embodiment of this invention a controllable solid propellant rocket engine capable of variable thrust during flight, the thrust control system being schematically depicted.

FIGURE 13 is a view corresponding generally to that of FIGURE 12 but showing an alternate control system utilizing electrical control means.

FIGURE 14A is a fragmentary view in longitudinal section of a solid propellant rocket engine, embodying means for elfecting safe non-propulsive storage in accordance with this invention.

FIGURE 14B is a view similar to that of FIGURE 14A but illustrating the use of a fluid spring for controlling the variable area nozzle and for rendering it non-propulsive during storage.

FIGURE 14C is a fragmentary part sectional plan view of a detail of FIGURE 14B.

FIGURE 15 is a block digram illustrating the sequence of operation for destruction of a disposable booster case employed in a solid propellant rocket engine embodying a variable area plug nozzle of this invention.

FIGURE '16 is a schematic diagram for an electrical control system for a disposable booster in the system of FIGURE 15.

FIGURE 17 illustrates an idealized pressure-time curve for the disposable booster in the system of FIGURE v1'5.

FIGURE 18 is a fragmentary view in longitudinal section through a conical shock diifuser having, in accordance with this invention, variable geometry in both rotation and translation.

FIGURE 19 is a fragmentary view in longitudinal section through a flexible bellows assembly of the diffuser of FIGURE 18.

FIGURE 20 is a plan view of an aircraft and/or missile having improved air breatihing engines embodying variable geometry fluid inlets and outlets in accordance with this invention.

FIGURE 21 illustrates the aircraft and/or missile of FIGURE 20 in a climb at an angle of attack, a.

FIGURE 22 is a fragmentary view in longitudinal sec tion of a further modification of a variable throat area nozzle in accordance with this invention, employing internal supersonic expansion based upon the principle of Pr-andtl-Meyer flow around a corner.

Nozzle With External Expansion FIGURES 1A, B and C illustrate the basic principle of a known, new type of nozzle which replaces the conventional converging-diverging or DeLaval nozzle. This nozzle, sometimes called a plug nozzle, is based upon the principle of Prandtl-Meyer supersonic flow around a corner. 7 It is essentially a reversal of the spike or Oswa titsch diffuser developed for supersonic inlets, the working fluid here undergoing external expansion rather than external compression as in supersonic inlets. The theory of this type of supersonic inlets and exits is reported in the literature and is described briefly with the aid of FIGURES 1A, B and C as follows. In FIGURE 1A, the plug nozzle consists of a central body, 26, which reduces in cross-section to form a spike or tip 26 having a streamline isentropic surface, 21, at its aft end. An annular passage, 22., is formed between the central body 20 and the outer shell, 23, which is curved inwardly at its aft portion ending in a short conical section at 24 which, together with the plug surface form a pre-determined cone angle, 9, for fluid flow, relative to the nozzle axis of symmetry, 8-8. The end of the nozzle shell at 25 is termed the lip, and the normal distance from the lip to the surface, 21, of the central body establishes the throat area (A,) of the nozzle. The nozzle shell, or outer structure is terminated at the lip, 25, although some internal supersonic expansion may also be provided in some cases as described later.

The working fluid from any source is conducted subsonically through the annular chamber, 22, at some high pressure relative to the ambient atmospheric pressure at the nozzle discharge, and is then directed radially inwardly at the predetermined angle, 0, formed by the nozzle structure. Sonic velocity is attained in the annular throat area at the lip 25. The working fluid then expands isentr'opically from Mach one at the lip to supersonic velocity as the flow turns the corner at the lip, forming its own free stream outer boundary, 32, with the atmosphere.

. The angle turnedby the fluid at the lip corner is a function of the pressure ratio or final Mach number and is denoted by the symbol, 1 referred to herein as the Prandtl-Meyer angle. For operation at the design pressure ratio, the angle 6 of the nozzle structure may be designed to equal the Prandtl-Meyer angle, 1/, so that the free stream jet boundary is then parallel to the axis of symmetry as indicated on FIGURE 1A. The exit area may then be defined by the lip diameter.

The streamline surface 21 of the central body preferably conforms to a streamline of flow under design conditions. The surface may be determined by the method of characteristics as described in any recent textbook on supersonic compressible fluid flow such, for example, as Elements of Aerodynamics of Supersonic Flows, by Antonio Ferri, 1949, The MacMillan Company, New York. When it is so formed, the internal shock losses within the jet are cancelled by superposition of expansion waves from the lip on compression waves reflected from the plug surface. When the surface is designed in this manner, it is referred to as an isentropic surface.

The isentropic surface, 21, FIGURE 1A, may be replaced by a simpler convergent conical surface, 21b, as shown in FIGURE LB, with only a moderate loss in thrust coefficient at lower design pressure ratios; or it may be replaced by a multiple conical surfaced plug with two or more conical angles, as illustrated in FIGURE 10, for somewhat higher pressure ratios, in which the shock losses occurring with a single cone are reduced. The isentropic surface of FIGURE 1A then represents the limiting case where the cone angle is infinitely variable for complete isentropic expansion. The exact shape of the conical plug may be selected to meet individual engine requirements as to design pressure ratios, desired nozzle efiiciencies, etc., the important feature being that external expansion is employed in preference to the conventional converging-diverging nozzle.

It is well known that plug nozzles provide substantially the same nozzle thrust coefficient as the convergent-divergent nozzles at the design pressure ratio. They offer superior performance at oil design pressure ratios lower than design value. External expansion results only in expansion to the local ambient atmospheric pressure, whereas convergentadivergent nozzles of fixed area ratio overexpand to sub-atmospheric pressures at lower off-design pressure ratios, reducing net thrust.

Nozzle With Variable T hroaz Area central body is made in the form of a hollow cylinder,

26, with radial slots or flow passages, 27, formed in a reinforced section 27', to which is rigidly attached a tapered plug 26' which, as depicted, has an isentropic surface 21, the continuity of structure providing minimum weight while carrying the large axial forces due to pressure acting internally on the plug. An annular cowl 28 is suitably, slidably, co-axially mounted on the cylinder 26 for axial movement relative thereto and suitably sealed thereto at its forward or left-hand end by an O-ring 37 in a groove 38.

The working fluid from any source (not shown) passes from the cylinder 26 through the radial slots 27 into the semi-toroidal annulus 34 and is thence directed aft, that is, to the right in FIGURE 2, between the cowl and plug. The normal distance from the cowl lip 25 to the plug 21 determines the throat area (A It can be seen that when the cowl is axially displaced aft to a position such, for example, as shown by the dotted lines at 28, the throat area increases, while a forward displacement of the cowl would reduce the throat area.

In FIGURE 2 the summation of working fluid pressure forces, indicated by the arrows, p, acting on the cowl will have a horizontal component acting aft. The pressure in the annulus above and to the left of the line AA will be substantially constant in all directions. In the converging channel between the cowl 28 and plug 21, below the line AA, the pressure will decrease, as indicated by the progressively shorter arrows, to the critical pressure at the throat. This pressure distribution acting upon the cowl portion near the lip will then produce an unbalanced force in the aft direction.

By slight modification as shown in FIGURE 3, the unbalanced pressure forces may be reduced or reversed in direction. This is accomplished by reducing the diameter at which the cowl 28 engages the cylinder 26 with respect to the maximum diameter of the plug, thereby increasing the area in the forward direction of the annulus 34. The unbalanced pressure force may then be reversed acting in the forward direction. Obviously, a diameter relationship may be used to produce a resultant control force approaching zero, or acting in either direction.

In FIGURE 2, the resultant pressure forces acting aft (10 the right) on the cowl 28 are transmitted by the rod 54, adjustable nut, 58, and compression spring, 57, sup ported by the brackets 55 and 56 mounted on the case 26.

In FIGURE 3, the resultant pressure forces acting forward (to the left), on the cowl 28, is resisted by the compression spring 57' and bracket 56' mounted on the case 26.

The control forces to position the cowl may be supplied in any desired manner. The simple mechanical springs in FIGURES 2 and 3 may be replaced by electrical, hydraulic or pneumatic actuators, with signals from the guidance or control system to vary the nozzle position in any prescribed manner, as described later.

T lzrust Vector Directional Control The preceding description explains the improvements obtained by translatory motion of the cowl to provide a variable throat area.

The complex problem of providing thrust vector directional control may now be solved in accordance with this invention by providing oscillatory motion of the cowl. This is illustrated in FIGURE 4 where the cowl 28 is pivotally mounted through a gimbal ring 41 to the thrust chamber structure 26. A spherical surface 36 is provided on the cylinder 26, which engages a similar surface on the cowl 28. An O-ring 37 in a groove 38 in the cowl 28 form a fluid seal, with a second elastomeric seal, 39, attached to the inner cowl 28 to protect the spherical surface 36. The cowl is pivotally attached at 40 to the gimbal ring 41, which in turn is pivotally attached at 42, from the pivot 40 to the support 43 on the cylinder 26. The cowl 28 is thus capable of being angularly dis- Z placed about a point on the centerline of cylinder 26 in the plane of the gimbal ring 431. This point is also the center of the spherical radius of the surface 36.

Since the flow turns through the same Prandtl-Meyer angle about the cowl lip, 25, as a focal point, rotation of the cowl causes rotation of the entire jet stream leaving the lip. The solid lines show the cowl 28 in normal or zero degree position, the line 32 being the free stream jet boundary with symmetrical flow about the centerline causing the thrust vector to act along the centerline also and, for the purposes hereof, serving also as zero reference line.

When the cowl is rotated clockwise as viewed in FIG- URE 4, say, about the pivot 42 to the broken line position shown at 28, the free stream boundary is also deflected in the clockwise direction as shown at 32'.

Similarly, when the cow-l is rotated, say 5 counter-clockwise to the broken line position shown at 28", the free stream boundary is also deflected in the counter-clockwise direction, as shown at 32". Means are provided for angularly displacing the cowl 23 on the pivots 42. Such means may be hydraulic, pneumatic or electric actuating means. As shown, an hydraulic cylinder 64 is pivoted to the case 26 and has a piston rod 64' pivoted to the cowl as at =64". As the piston is moved axially by the cylinder, the cowl will be caused to tilt on the pivots 42. In a similar manner, the cowl may be rotated in a plane normal to the paper, or about the pivot 40 through the provision of other hydraulic cylinders 6'4, 90 from the pivot 40; or about both planes simultaneously as is well known in gimbal mounted structures.

The angular rotation of the thrust vector will then be very nearly equal to the angular rotation of the cowl. Lossess will necessarily occur due to the assymetric flow about the spike when the cowl is displaced from normal position. Since the required displacement is only a few degrees, the losses are not severe and occur only during control. a Thus, it may be seen that thrust vector directional control may be provided within the nozzle itself, eliminating the weight, drag, and cost of additional subsidiary means previously described, such as jet vanes, or jetavators. This then eliminates the need of additional high temperature resistant components, the nozzle itself serving this function, thereby improving reliability.

Since pressure balancing the cowl reduces the control forces required to position it, this also greatly reduces the forces acting on the gimbal ring structure supporting the cowl. The assembly may then be designed so that the loads on the gimbal ring structure are only a small fraction of the total thrust load which must be carried by the gimbal structure when the entire thrust chamber assembly is gimbaled, as in liquid propellant rocket engines.

The moment of inertia of the cowl is substantially reduced, compared to rocket thrust chamber assemblies, so that the actuating forces may be reduced with faster response provided by the system shown here. The thrust chamber may now be rigidly mounted to the missile, eliminating need of flexible feed lines, again reducing weight and improving reliability.

Thrust Termination Means for thrust termination is a major problem with solid propellant rocket engines. It is known from the internal ballistics of solid propellant rocket engines that, for any given grain design, the chamber pressure decreases as the throat area is increased. By means of the variable throat area nozzle described previously, the throat area may be increased continuously, with a continuous decrease in chamber pressure, thereby avoiding the abrupt discontinuity which occurs with nozzle insert ejection, previously described. In this manner, thrust decay can be reproducibly controlled simply by extending the cowl 28 so as to increase the throat area to many times its normal design value. When the chamber pressure reduces below the critical pressure ratio, even if combustion of the propellant grain continues, the working fluid would have a random velocity distribution at the exit, rendering the unit non-propulsive. With propellant properties which permit continued burning at low ambient atmospheric pressures, the cowl may again be retracted until at super-critical pressure ratios, the thermal energy of the gases is again converted into kinetic energy of the jet stream, with high velocities directed aft, again rendering the unit propulsive.

The above system has the advantage that Vernier control over both missile velocity and attitude may be provided by operating at reduced thrust levels with vector control available before thrust termination.

For installations where variable area nozzles are not desired, or, in the event faster response is desired, thrust termination may be provided by only slight modification of the nozzle structure, as shown in FIGURE 5A. In this method, the cowl 28 is fabricated in longitudinally split sections joined by electrically fired explosive bolts, 50. When it is desired to terminate thrust, a signal detonates the explosive bolts, the cowl thereby being released from the cylinder 25 under the action of the internal gas pressure within the cowl annulus. The large increase in flow area is chiefly in a non-propulsive direction, thereby reducing peak thrust loads associated with ejection of a nozzle insert. Even if propellant continues to burn at reduced chamber pressures, the gases are discharged radially through the ports 27, producing no thrust.

Variation of this principle is illustrated in FIGURE 5B in an improved form. The cowl 28 maintains its integral structure, but contains large radial ports, 51, normally closed by the retainer ring 52 protected by thermal insulation 53. The retainer ring is held in closed position by the explosive bolts 5b. Upon signal, detonation of the bolts releases the retainer ring and insulation, the increased flow area reducing chamber pressure as above. The annular throat area between cowl Z8 and surface 21 now remains intact.

Thus, it may be seen that improved means for thrust termination may be provided which eliminate thrust peaks as with ejection of nozzle inserts; simplifies problems relating to disposition of gases expelled in a forward direction; provides no interference with internal ballistics, or grain installation, and provides minimum weight.

Combined Variable Area, Thrust Vector Directional Control, and Thrust Termination Since thrust termination is associated chiefly with details of cowl construction, it is readily applicable in combination with other improvements.

The combination of variable area and thrust vector control requires means to provide both 'translatory motion and oscillatory motion of the cowl. This is shown in FIGURE 6, in which the cowl 28 embodies thrust termination ring 52 and in which the same numerals again refer to similar parts.

A bellows seal, 59, replaces the spherical surface and O-ring seals of FIGURE 4, providing convenient means for obtaining the desired freedom of motion. Oscillatory motion may again be provided by a gimbal structure generally similar to FIGURE 4. The rigid support, 43, of FIGURE 4 for attaching the gimbal ring 41 to the cylinder 26 in diametrically opposite sides of the cylinder is now replaced by new means which allow the gimbal ring to move axially with respect to the cylinder. The pivot 42 is now attached to a shaft, 54, slidably mounted on journals 55 and 56 fixed on the cylinder 26. Such shaft 54 is positioned by the spring 57 through an adjustable nut 58 which adjusts the spring force to provide the desired control. Obviously the spring may be replaced by a fluid or electrical actuator, conforming to the missile control system, as desired. Variable area is then provided by compression or extension of the spring 57 causing translation of the gimbal ring and cowl assembly.

Linear actuators, such as the hydraulic cylinder 64, are pivoted to the case 26, with piston rod 64', pivoted to the cowl 28 at 64", to provide rotation of the cowl 28 for thrust vector directional control, in the manner previously described.

An alternate seal to the bellows design, but still providing the freedom of motion, is shown in FIGURE 7. This is a flexible seal, 70, similar to a diaphragm or boot, but better adapted to carry pressure forces, and is described more fully in my co-pending patent application Serial Number 432,745, now abandoned, and decribed briefly as follows: one end, 71, is rigidly secured to the case 26, with a loop 72 formed between this and the opposite end 73 rigidly attached to the cowl 28. As the cowl is displaced axially, (translatory motion) relative to the cylinder 26, the flexible seal rolls from one surface to the other. As the cowl oscillates, the seal conforms to the new wall angle of the cowl in one plane, while it undergoes a few degrees of torsional deflection in a plane 90 to the first plane. Since the sleeve is fabricated of reinforced elastomeric materials, the various deflections and radial elongations are easily provided. A second elastomeric seal, or trap, 74, attached to the cowl, or case, or both, may be provided similar to a labyrinth seal. This would entrap stagnant gases in the region of the flexible seal or bellows thereby reducing the heat transfer rate. The seal may be fabricate-d of greater thickness than normally required to allow for some deterioration of the surface in contact with high temperature gases.

An alternate method for providing both translatory and oscillatory motion of the cowl may be provided without gimbal rings by use of four actuators mounted 90 apart, as shown in FIGURE 8. The cowl 28 is sealed to the cylinder by a bellows 59 or flexible seal as previously described. The four actuators, 203, which may be linear electrical actuators with self-locking screwthreads, are located 90 from each other and are pivotally mounted at their forward end to the cylinder 26 by the pin 204 through the bracket 205 fixed on the cylinder 26. The aft end of the actuator rod, 208, passes through a cross-slide, 2%, carried by the crossways, 207, so as to provide freedom for the actuator to rotate in a radial plane, but which restrains it in a circumferential direction, as illustrated in FIGURE 8B, taken on line line 313-83 of FIGURE 8.

The actuator is attached to a suitable lug, 209, on the cowl 28 by means of spherical ball joint or rod end, 210, providing freedom of relative motion in two planes; or rubber bushings may be used which accomplish the same purpose for small angular deflections.

The four actuators may then operate simultaneously uni-directionally to either extend or retract the cowl to achieve variable area. In order to provide vector control, any two diametrically opposed actuators operate contradirectionally, while the remaining two actuators remain fixed in locked position. The latter two actuators establish the center of rotation for the opposite pair.

Combined Variable Area, Thrust Vector Control and Thrust Termination with Internal Supersonic Expansion The improvements shown herein are not necessarily limited to nozzles with external expansion. It is possible to combine variable area, thrust vector control and thrust termination within a nozzle employing only internal expansion. This is shown in FIGURE 9 in which the plug, 211, is made integral with the engine thrust chamber, tailpipe, or rocket case, as represented by 26, while the cowl 28 includes a diverging portion of the nozzle, 2 12, and is flexibly mounted as by means of bellows 59, or, alternately, as by seal 70 shown in detail in FIGURE 7. Again, the throat area may be varied by translatory motion of the cowl effected by means of the cowl actuators, 203, which also varies the expansion ratio, while thrust vector directional control is provided by rotation of the cowl and diverging cone, in a manner previously described. Thrust termination is also provided by releasing the retaining ring, 52, in the cowl, also as already described.

Application of these new principles to various types of jet engines and their resulting improvements may now be described.

Improvements in Solid Propellant Rocket Engines A variable throat area nozzle offers important mechanical solution to many problems in solid propellant rocket engines. It makes possible substantial reduction in case weight beyond the use of higher strength materials by automatically maintaining constant chamber pressure independent of the temperature sensitivity of the propellant, and of the progressivity of regressivity of the propellant grain.

It also makes possible a controllable solid propellant whose thrust may be varied at will, providing flexibility in operation which even surpasses throttleability of liquid propellant rocket engines.

FIGURE 10 illustrates the application of the new nozzle principles to a solid propellant rocket engine, in which the same numbers are used to identify parts previously described. A solid propellant grain, 215, is enclosed within the case, 26, to which is rigidly attached a plug 26' having an expansion surface, 21, which may be a cone, double cone, or isentropic surface. A cowl, 28, attached to the case is nearly pressure balanced, having a differential pressure force acting aft. This force is balanced at diametrically opposite points by the simple mechan ical spring, 57, through the shaft, 54, and nut, 58, in a manner similar to that described for FIGURE 6. The adjustable nut 58 provides initial compression of the spring 57. This retains the cowl in its completely retracted position during storage, the gasket, 99, preferably of a resilient elastomeric material providing a vaportight seal.

A bellows seal 5Q, such as is shown in FIGURES 6 and 8, is preferred between case and cowl, even when vector control is not required; this provides freedom of motion, eliminating costly sliding surfaces requiring close tolerances on diameter and roundness, reduces friction and eliminates binding due to unequal thermal expansion or deposit of foreign particles on the sliding surfaces. The cowl is maintained concentric automatically by the internal pressure forces when in operation.

Upon ignition (by conventional means not shown), the internal gas pressure generated by combustion acts upon the unbalanced area of the cowl. As the pressure force on the cowl overcomes the initial spring force, the cowl opens ejecting the gasket 99 and compressing the spring 57 until the spring force balances the pressure force. The latter may be made relatively light depending upon the amount of area unbalance designed into the cowl, as already previously described, so that relatively light springs may be used.

It may be seen that when the internal pressure increases (due to any of several reasons as described below), the higher pressure force causes further compression of the spring 57, opening the cowl to a larger throat area. Similarly, when the internal pressure decreases, the lower pressure force causes the spring 57 to extend, thereby retracting the cowl, resulting in a smaller throat area. The throat area thus increases automatically with increasing chamber pressure.

Application of the concept of variable throat area to the internal ballistic equations for solid propellant rocket engines reveals that as the throat area increases, pressure will decrease by some higher power; thrust and mass flow rate will both decrease. The latter two decrease with increasing throat area since the pressure reduces faster than the area increases.

The inverse exponential relationship between the parameters pressure, thrust and mass flow rate as func- Constant Pressure Operation Independent of Ambient Temperature Solid propellants are temperature sensitive, the propellant burning rate increasing as the temperature of the grain, before firing, increases. The increased burning rate results in higher rate of gas formation, increasing chamber pressure and thrust with constant throat area nozzles. Conversely, burning rate, pressure and thrust decrease at lower ambient temperatures.

As an example, if the rocket is designed to operate at 1,000 p.s.i. at 60 F., the temperature sensitivity of one type of propellant may increase the operating pressure to more than 1200 p.s.i. at 160 F, with a constant throat nozzle, or by 20%. Similarly, if the rocket is fired at extremely low ambient temperatures, at say 40 F. (100 below ambient), the chamber pressure reduces to approximately 800 psi. or by minus 20%.

The rocket must be designed to provide the desired thrust at its minimum operating temperature, but the case must be designed to withstand its maximum operating pressure at the high temperature which is 50% higher pressure than the minimum, resulting in approximately a 50% weight penalty for the rocket case.

The eiiect of the variable area nozzle in providing a substantially constant pressure is illustrated by the curves of FIGURE 11, in which chamber pressure is plotted against throat area. The curve A-A may rep-resent the change in chamber pressure with increasing throat area for a neutral burning propellant grain at ambient temperature of 60 F. The curve B-B represents the chamber pressure at the maximum ambient temperature of 160 F., While curve C-C represents the same quantity at a minimum ambient temperature of -40 F. The abscissa A represents the normal throat area for the variable area nozzle and also the value of an equivalent conventional constant throat area nozzle. The ordinate at the normal throat area intersects the curve AA at the normal design pressure P (say, 1000 psi). If the nozzle area remained constant, the chamber pressure at the higher temperature corresponds to the intersection with the curve BB, the maximum pressure increasing from P to P (or to 1200 p.s.i.). Similarly, at the minimum temperature represented by curve C-C, the chamber pressure decreases to the value P (800 p.-s.i.).

With the variable area nozzle, the spring 57 provides the characteristic indicated by the line D--D of FIG- URE 11, where the intersection of this line with the P axis represents the initial compression F in the spring (obtained by adjusting the nut 58). The slope of the line DD represents the increase in spring force resulting from further compression of the spring which has a spring rate of k pounds per inch of deflection.

As the pressure on the nozzle increases at the elevated temperature, the nozzle opens beyond its normal position, the increased throat area reducing chamber pressure until the spring force and pressure force come into equilibrium where the line D-D intersects the curve B-B. The corresponding chamber pressure is then only P only slightly higher than the normal pressure P the increase being due only to the spring rate of the spring 57. Use of low rate springs obviously makes this increase low, compared to P corresponding to the fixed throat nozzle.

Similarly, at reduced temperature, the reduced chamher pressure causes the nozzle to retract until the line DD intersects the curve CC, the reduced chamber pressure P being only slightly below the normal pressure P Constant Pressure Operation Independent of Progressivity and Regressivity 0f the Propellant Grain When the burning surface of the propellant increases as burning progresses, the additional burning surf-ace causes a higher rate of gas formation, increasing chamber pressure and thrust. Such grains are described as progressive burning grains. Similarly, a reduction in burning surface reduces chamber pressure and thrust and is known as a regressive burning grain, While a constant burning surface is known as a neutral burning grain. The latter is required to provide constant thrust throughout the total burning time.

While great strides have been made in recent years in improving grain geometry, particularly with internal star configurations, some degree of progressivity and/or regressivity exists. The effect on chamber pressure and case Weight is identical to changes in ambient temperature. In this case, the curve A-A may represent the initial burning surface after ignition, While the curve B-B may represent the maximum burning surface of a progressively burning grain. It may be seen that with constant area nozzles peak design pressures occur with progressive burning grains which again penalize case Weight.

Again, the curve C-C represents minimum burning surface with a regressively burning nozzle. The automatic operation of the variable area nozzle in reducing pressure variation due to progressivity or regressivity will be the same as previously described for temperature variation. This offers greater degree of freedom in the design of propellant grains, since the pressure peaks and case weight penalties associated with a high progessivity ratio may now be minimized by the variable area nozzle. 7 The combined effect of grain progressivity ratio plus temperature sensitivity of pressure can lead to extreme peak pressures, where the use of the variable area nozzle principle is most helpful.

Reduced Thrust Variation in Solid Propellant Rocket Engines The variation in chamber pressure with ambient temperature of the propellant grain before firing causes even greater variation in rocket thrust when constant area nozzles are used, since the nozzle thrust coeflicient also increases with increased pressure.

In the example cited above, the thrust at P. will be more than 50% greater than the thrust at 40 F. This Wide variation in thrust is undesirable, aifecting missile performance, maximum accelerations, as well as penalizing missile Weight.

When the chamber pressure is maintained substantially constant over the ambient temperature range by means of the variable nozzle, the change in thrust will be due only to the change in throat area. Due to the inverse exponential relationship between chamber pressure and nozzle throat area, a much smaller change in throat area results in a larger change in pressure.

As an example for a typical burning rate exponent of n equal to 0.50, the pressure is inversely proportional to the second power of the throat area. A 50% pressure variation from 40 F. to 160 F. may then be restored to its normal value by only a 22% increase in throat area and, hence, in thrust. This is less than half the thrust variation for the constant area nozzle.

Propellants having higher values of the exponent n will produce even smaller thrust variation. If another propellant is used having a value of n equal to 0.75, the pressure is now inversely proportional to the fourth power of the throat area. A 50% pressure increase from 40 F. to 160 F. may now be compensated by a 10.7% increase 13 in throat area with the variable area nozzle, the thrust varying by less than one-fourth compared to the constant area nozzle.

Thrust variation at extreme temperatures about the normal ambient temperature of 60 F. is of interest. in the constant area nozzle, the thrust variation at each temperature extreme is more than plus or minus 20%; for the variable area nozzle, the variation is plus or minus 9.5% for 11:0.5; and plus or minus 4.7% for n=0.75.

This illustrates another important advantage of variable area nozzles. It makes propellants with higher values of n more desirable for rocket use. For constant area nozzles, high values of n" are undesirable because of the rapid increase in chamber pressure with small variations in burnin surface, nozzle throat, and temperature. For variable area nozzles, high values of 11 provide more uni'form,,thrust, requiring. smaller changes inthroat area to compensate for such effects.

The above description illustrates the eifect of a variable area nozzle in reducing thrust variation between rounds fired at various ambient temperatures. The same relationships also hold for reducing thrust variation occurring during a single firing as a result of grain progressivity or regres-sivity. It is seen that it is possible to produce nearly constant thrust (within plus or minus with a grain progressivity ratio of 1.20 while maintaining constant chamber pressure and, hence, minimum case weight, with variable area nozzle. As a further example, a progressivity ratio as high as 2.0 (160% increase in propellant burning surface) will increase thrust by less than 18% when n=0.75, while the variable area nozzle still maintains substantially constant pressure and minimum case weight.

Adjustable Thrust Solid Propellant Rocket Engine The same rocket engine, as described in FIGURE may be used at various thrust levels simply by adjusting the nut, 58, to provide different amounts of initial spring compression on the spring, 57.

For example, the initial spring compression, F may be decreased as represented by the line D'D' of FIG- URE 16; the reduced chamber pressure at normal ambient temperature would be represented by the intersection of DD with the line AA. The corresponding throat area and chamber pressure would then produce a lower thrust level, since the pressure reduces faster than the throat area increa es. The spring 57 would automatically maintain a nearly constant chamber pressure about the new pressure level, as previously described.

In a similar manner, the thrust level could be increased to a line above DD (not shown), the increase being limited by the maximum design pressure of the case.

The overall impulse to weight ratio would deteriorate as thrust was adjusted downwardly, since the specific impulse would decrease at lower pressures, and the case would necessarily be heavier than required at the lower thrust level.

Manual adjustment, as described, may also be made for extreme ambient temperature conditions, thus making thrust more nearly exactly reproducible at all temperature limits.

A Controllable Solid Propellant Rocket Engine The use of a variable area nozzle makes possible a controllable solid propellant rocket engine in which the thrust may be varied in magnitude at will throughout flight. Thrust variation may be obtained by varying the nozzle cowl position and, hence, throat area responsive to any type control system.

The mechanical springs for positioning the cowl as shown in FIGURE may be replaced by servo-controls, either fluid or electrically actuated, illustrated in FIG- URES l2 and 13 respectively.

In the systems shown in FIGURES l2 and 13, the difference between rocket chamber pressure and a control 1 1 pressure from any source is used as the input signal to increase or decrease thrust.

A system using fluid actuation is illustrated in FIGURE 12, which shows a fragmentary view of the solid propellant rocket engine of FIGURE 10, in which a fluid actuator, 1121, supported by the case 26 and attached to the cowl by the pin 102 is responsive to the control valve 103. The latter consists of a body 104 in which is positioned a sliding spool 1&5 containing ports 1% and 107, The body contains a high pressure inlet at 108 and two vent or return lines 199 and 11d. Displaced circumferentially are two additional ports, 111 and 112 joined by tubing 113 and 114 to opposite sides of the piston 115 within the actuator 161.

One end of the control valve 193 contains a bellows 116 communicating through tubing 117 with the rocket chamber. lows 116 against the spool 1125 and is opposed by a second bellows 113 within the valve body and joined by tubing 119 to receive a controlled pressure, as from a pressure regulator 120 from the missile hydraulic system. A three-way on-oii valve 121 including a vent 122. vents fluid from the bellows 118 through port 122 when the valve is closed.

In operation of this system, when the control pressure force from bellows 118 is balanced by chamber pressure force in bellows 116 the spool 165 is centralized in neutral position with no flow through the control valve 103. When the control pressure in bellows 113 is increased by the missile guidance system, the spool 165 is momentarily displaced to the right; high pressure fluid from inlet port 108 passes through ports 1% and 111 and through tubing 113 to the actuator 161 so as to retract the cowl. This increases rocket chamber pressure, and as its value approaches the control pressure, the bellows 116 returns the spool 105 to its central position closing ports 108 and 111. As described previously, retraction of the cowl increases chamber pressure by a greater amount than the reduction in throat area, increasing rocket thrust.

When the missile control system reduces the control pressure supplied to bellows 118, the above operation is reversed, causing the cowl to extend, reducing chamber pressure and thrust.

A system using electrical actuators is illustrated in FIGURE 13 operating on the same principles described above. The spool 165 between bellows 115 and 118 is replaced by the electric switch 144 comprising two conductors 145 and 146 separated by an insulator 147. A brush 148 connected to ground makes contact with the conductor 145 or 146, or the insulator 147. A linear electrical actuator consists of a reversible electric motor having split windings 149 and 1511 which drive the screw 151 linearly in a forward or aft direction. This screw may have a self-locking thread. One terminal of the split windings is connected to a common battery 152, or other source of electric energy. Wires 153 and 154 connect to opposite terminals of the split windings to the conductors 14-5 and 145 respectively. Limit switches 156 and 157 in the wire circuits 153 and 154 open the respective circuits when the screw thread reaches its limit of travel in either direction.

It can be seen that when the control pressure in bellows 118 is higher than the rocket chamber pressure, the conductor 145 contacts the brush 148 closing the circuit energizing coil 149 causing the actuator to retract the cowl, increasing chamber pressure and thrust. When the control pressure reduces to a value less than chamber pressure the conductor 1 .6 closes the circuit energizing the coil 150, causing the actuator to extend the cowl, reducing chamber pressure and thrust. When the pressures become equalized, the switch 144 reaches neutral position, the insulator 147 opening both circuits, maintaining cowl position and thrust.

With either fluid or electrical actuators, in a limiting case, the throat area may be increased so that the chamber Fluid pressure from the chamber acts within bele 15 pressure reduces below the minimum pressure at which the propellant will burn, thereby causing thrust termination. With propellants which continue to burn at low ambient pressures, the unit may become non-propulsive, as previously described and illustrated in FIGURE 8. In this case, re-starts are possible Without re-i-gnition, permitting a boost-glide thrust-time characteristic.

Thus, it may be seen that a controllable solid propellant rocket engine may have its thrust-time curve varied at will during any one firing, or from one firing to the next, even though the same grain is used, responsible to the input signal supplied to the servo-mechanism.

It may be noted that the servo mechanism, either fluid or electric, will still compensate automatically for variations in" chamber pressure due to temperature sensitivity of the propellant, and/or due to progressivity or regressivity of the propellant grain. It will automatically maintain constant pressure at any 'vflue set by the control pressure.

The actuators, either fluid or electrical, of the servomechanisms may be used for both thrust vector directional control as well as thrust magnitude control, as de scribed previously in connection with FIGURE 8.

Non-Propulsive Storage for Solid Propellmzt Rocket Engines Means for rendering solid propellant rocket engines non-propulsive during storage are sometimes desired. This avoids catastrophic damage in the event of accidental ignition of the propellant grain.

Solid propellant rocket engines equipped with means for thrust termination, as described and shown in FIG- URE B, may be readily stored in non-propulsive condition simply by removing the retainer ring, 52. Fast acting self-locking clamps may be used (not shown), to assemble the retainer ring just prior to firing as an arm operation.

Not all solid propellant rocket engines require thrust termination, in which case, other means are required to render the unit non-propulsive during storage.

For the controllable solid propellant rocket engine using fluid actuators illustrated in FIGURE 12, non-propulsive means may be provided by the normally closed three-way on-oii valve 121 with vent 122. When closed, the bellows 118 cannot be pressurized, since even in the event of leak-age past valve 121, the vent 122 prevents pressurization. In the event of accidental ignition, chamber pressure in the bellows llo'will displace the spool 105 in control valve to the left, the return line 113 from actuator 161 being then connected through port 195 to vent 1111. The unbalanced pressure forces on cowl '28 will then displace fluid from actuator 1111 through tubing 113 permitting the cowl to extend, rendering the unit nonpropulsive. If desired, a second valve 123, normally open to vent 124, may be connected to line 113 to accomplish the same thing. This would require that valve 123 be positively closed as an arming operation prior to firing.

For the controllable solid propellant rocket engine using electrical actuators, illustrated in FIGURE 13, non-propulsive storage may be provided by full nozzle extension during storage. Accidental ignition would produce no thrust at sub-critical pressures with large nozzle throat area.

The master control switch 158 of FIGURE 13 would be cloesd as an arming operation prior to firing, the on-off valve 121 being energized on arming, admitting fluid to bellows 113, thereby closing circuit 154 and retracting the nozzle cowl to its minimum throat area position. When the igniter is energized by the Fire switch, the reduced throat area renders the rocket propulsive, the chamber pressure now rising and being controlled by the differential pressure switch 144 of FIGURE 13.

For the constant pressure solid propellant rocket engine using mechanical springs to position the cowl, as shown in FIGURES and 6, non-propulsive storage may be provided by rendering the springs inoperative during storage until armed prior to firing. This is illustrated in FIGURE 14A where the shaft 54 attached to the cowl is restrained by the spring 57 only when the clutch assembly 131 has been engaged to move with the shaft 54. During storage, the pin 132 is maintained in a disengaged position by the spring 133. In the event of accidental ignition, the shaft 54 is free to move independently of the spring 57, the cowl advancing to its fully extended position, fixed by the stop 135 on the forward end of shaft 54. A light coil spring 134 may be used optionally to keep the cowl positioned during storage and shipment, but offers slight resistance to motion under the unbalanced pressure forces on the cowl.

The engine would be armed prior to firing by the fluid actuator 136 (or by a solenoid actuator, not shown), causing the pin 132 to engage the shaft 54 through the opening 138/ In this condition the clutch assembly 131 is locked to the shaft 54, engaging the spring retainer 139 and compressing the spring 57, the spring and cowl pressure forces then coming into equilibrium as previously described.

For convenience in assembly, the spring 57 is precompressed and calibrated within the housing 140, the retainer 139 and adjustable nut 53, the latter being locked in position by the set screw 141.

The mechanical springs 57 positioning the cowl may be replaced by fluid springs, the springs being made inoperative by deflation fcr non-propulsive storage, and operative by inflation to the required pressure to provide the desired spring characteristics to render the engine propulsive.

-A fluid spring installation is illustrated in FIGURE 14B and 14C in which the fluid spring assembly 159 replaces the mechanical spring 57 and associated parts shown in FIGURE 14A. The outer cylinder 16%; is supported by the bracket 161 attached to the case 26. A yoke 162 attached to the cowl 23 transmits force on the cowl to the inner cylinder 163 of tha -fluid spring. A flexible seal 164- similar to that previously described in FIGURE 7 provides a ileakpro of seal between inner and outer cylinders. A solenoid operated three-way valve 165 having a high pressure fluid inlet line 16 6, an outlet line 167 connected to the fluid spring, and a vent port 168 inflates or deflates the spring as required. The fluid inlet line 166 is connected to a pressure regulator 121 to which fluid is supplied from an accumulator 169 containing a compressible fluid under high pressure.

During storage, the fluid spring may be vented through the valve 168, thus offering no resistance to motion of the cowl 28. In the event of accidental ignition, the nozzle throat area may increase rendering the rocket nonpropul-sive. When the rocket is armed, the valve 165 is energized, closing vent 168 and supplying regulated fluid pressure through valve 165 to the fluid spring. As the spring force increases under fluid pressure the spring extends until the nozzle cowl retracts to its minimum throat area. The inflation pressure in this position determines the initial spring force F and the spring characteristics as illustrated by lines DD or DD' of FIGURE 11, thus establishing the thrust level of operation.

For some applications, the accumulator 169 and pressure regulator 129 may be part of the ground launching equipment, the line 166 being an umbilical cord connection having a quick disconnect and check valve.

Combined Non-Propulsive Storage and Thrust Termination The means shown in FIGURES 14A, B and C for rendering the rocket non-propulsive may also be used for controlled thrust termination during flight. With the constant pressure, solid propellant rocket engine which this system provides, only two cowl positions are involved: its normal propulsive position automatically con- 17 trolled when the springs are operative; and its fully extended position for thrust termination.

This may be readily accomplished by the fluid spring system shown in FIGURES 14B and 14C, simply by returning the valve 165 to vent position, releasing fluid pressure in the fluid spring assembly, thereby allowing the cowl to extend to its maximum throat area position.

For the mechanical spring system of FIGURE 14A, only slight modification is required. Means are required for introducing high pressure fluid below the piston head of the pin 132 to disengage the clutch assembly from the shaft 54. One method would be to energize the squib 143 pro-assembled within the fluid actuator 136. Alternate means would introduce high pressure fluid below the Disposable Booster Solid propellant rocket engines are widely used as booster rockets to launch groundetmair missiles for defense against aerial attack. When the booster rocket burns out, it separates from the missile and falls freely to Earth. When such missiles are used to defend populated areas, the falling booster represents a hazard to the friendly civilian population. Means for disposing of the booster in the air after burnout is highly desired.

The variable area nozzle ollers means for disposing of the empty booster case by fragmentation in the air after separation by increasing chamber pressure to explosive force. This maybe accomplished by the following technique.

Controlled thrust termination is provided by increasing nozzle throat area before propellant exhaustion, conserving a small amount of propellant, probably equivalent to the remaining slivers, or slightly larger. Thrust termination permits separation to proceed in the conventional manner. After separation, the nozzle throat area is again reduced to a minimum value approaching zero. The remaining propellant is i e-ignited and due to the small throat, the pressure time curve increases asymptotically. With light weight cases, fabricated of highly heat treated steel, brittle fracture would result due to the high rate of loading. If desired, the case may be scored in any prescribed manner, introducing notch sensitivity at critical masses, reducing individual fragments below lethal size. The closed plug nozzle, which acts as a pressure vessel closure, is also more easily rupturable, in contrast with the heavier conventional open DeLaval nozzle which is lightly stressed.

The above sequence is shown schematically in the block diagram of FIGURE 15. While any of the variable area nozzle cont-rol systems may be adapted to operate in this sequence, the electrical system shown in FIGURE 16 represents one method for adapting the constant chamber pressure fluid spring system, illustrated in FIGURES 14B and 14 C, to operate as a disposable booster also. The additional equipment required to accomplish this are: (1) two time delay switches; (2) a secondary igniter; and (3) a squib assembled within the fluid spring, or alternate means to increase fluid spring pressure.

Referring to FIGURES 15, 16 and 14B and C, closure of the Arm switch 170 energizes the solenoid valve 165 through the normally closed primary time delay switch 171, thus inflating the fluid spring assembly 159, rendering the spring operative and the rocket propulsive. Electrical energy may be supplied from any source, such as the battery 172. A single wire system is illustrated, all components having a common ground.

Closure of the Fire switch 173 energizes the primary igniter 174 initiating thrust and simultaneously initiating the time cycle in the primary time delay switch 171. This switch may be of any type, such as an electrical heating element for diflerential thermal expansion of a bi-rnetallic strip supporting an electrical contact. The primary time delay switch controls thrust duration,

which ends after seconds, the time constant for the primary switch, which then opens the circuit de-energizing the solenoid valve venting high pressure fluid from the spring. The cowl is then free to extend, the increased throat area reducing chamber pressure .and causing thrust termination. The booster then separates from the missile in its conventional manner as with propellant exhaustion.

Physical separation of the booster from the missile causes the micro-switch to close contacts 176 and 17 7, the first contact 176 re-energizing the solenoid valve, closing its vent; the second contact 177 activating the secondary time delay switch 178 and the squib 179. The latter is assembled within the fluid spring and generates fluid at high pressure to retract the cowl and maintain it in closed position, the fluid spring pressure and, hence, spring force being substantially higher than the initial inflation pressure controlling thrust. Alternately, the squib 179 may be replaced by a solenoid valve which supplies fluid to the spring at very high pressure, such as directly from the accumulator 169 or by readjustment of the pressure regulator 120.

The secondary time delay switch 178 provides time for the missile to continue on course out of range of possible damage from booster fragmentation. This switch is normally open, and after t seconds, it closes, energizing the secondary igniter 180. This re-ignites the remaining propellant charge, but since little gas can escape due to the small throat area (high value of K), the pressure rises to explosive force within the rocket case. A typical booster chamber pressure-time curve is illustrated in FIGURE 17.

In the event a propellant is used which continues to burn at the low ambient pressures after separation, the second igniter 180 may be eliminated. In that event, the secondary time delay switch 178 would then energize the squib 179 or alternate means after t seconds, the high spring forces causing rapid nozzle closure and producing the same result.

The mechanical spring system of FIGURE 14A may be adapted to perform in a similar manner. Arm and thrust termination may be provided by a solenoid 4- Way valve replacing the solenoid valve 165 of FIGURE 14B and FIGURE 16. The electrical circuit as shown operating the 4-way valve alternately to supply high pressure fluid first above the piston head of the pin 132 for arming, and at expiration of t seconds, de-energizing the solenoid to reverse the flow, supplying fluid below the piston head of pin 132, dis-engaging the mechanical spring from the shaft 54 for thrust termination. A separate fluid actuator (not shown), would be required containing the squib 179 in order to retract the cowl. It may be located in tandem at forward end of shaft 54, or displaced circumferentially and attached to the cowl.

Both the fluid and electrically actuated controllable solid propellant rocket engines shown in FIGURES l2 and 13 may also be adapted to destroy the rocket in flight. Means for arming and thrust termination would be the same as already described for arm and non-propulsive storage. The added requirement of returning the cowl to closed position may be accomplished in a manner similar to that described.

While means responsive to physical separation of the booster from the missile have been described for initiating the remaining cycle for disposing of the booster, other means may be used, such as pressure switches responsive to chamber pressure decay on thrust termination; limit switches responsive to cowl position, or other rneans. Also, the propellant may be allowed to burn to exhaustion without time delay switch to signal thrust termination. Destruction of the case may then be achieved by detonation of a high explosive charge stored within and insulated from the propellant.

It is believed the system described is safest and most positive. Malfunction of any of the added components I required for booster disposal will not interfere, in most cases, with normal effectiveness of the missile, but would be limited only to failure to destroy the booster.

Improvement in Air Breathing Engines llmproved performance of air-breathing engines may be obtained not only by the use of improved discharge nozzles, substantially as described, but also by application of similar principles to improve the supersonic inlets, or diffusers.

The advantages of variable throat area nozzle for turbojet engines is well known, with considerable work having been done in variable area nozzles of the iris-type. This work has been limited to variable throat areas whereas the improved nozzles described herein when applied to turbo-jet or ram-jet engines provide variable area expansion ratio and thrust vector directional control as well. The advantage of variable expansion ratio has been described. The use of vector control with air-breathing engines is new.

With vector control, the aerodynamic control surfaces on aircraft or missiles may be reduced in size, or eliminated. Greater maneuverability may then be obtained in flight at great altitudes in rarified atmospheres by use of a component of engine thrust to produce forces normal to the longitudinal axis of the vehicle. A nozzle for turbo-jet or ram-jet engines which makes possible both vector control as well as variable areas then offers distinct advantages.

With improved exit nozzles, performance is still limited by inlet conditions. Fixed geometry inlets provide optimum performance at only the design flight Mach number, with thrust and specific fuel consumption suffering at speeds below and above the design value. In the case of ram-jet propelled missiles, this limits its ability to accelerate under its own power, requiring larger booster rockets to achieve higher boost velocities before the ramjet missile becomes operational.

Rate of climb of supersonic vehicles powered with airbreathing engines is limited because of loss of thrust at higher angles of attack due to asymmetric flow at the supersonic inlet.

' Improvement in aircraft and missile performance in the supersonic regime may be achieved by means of the flexible inlets or diffusers described below.

Flexible Supersonic Inlets Ram-jet and turbo-jet engines operating at supersonic flight speeds frequently employ conical shock diffusers, sometimes called spike or Oswatitsch type diff-users, in which external compression is obtained by oblique shock waves originating at the tip of the cone, or inner body, and which meet the cowl lip at the design flight Mach number.

This type diffuser has been extensively described in the literature on supersonic compressible fluid flow; these references also describe the improvements possible by variable [geometry inlets in which the position of the spike or inner body is displaced axially relative to the cowl in order to maintain the oblique shock on or near the cowl lip at off-design flight Mach numbers. This provides a variable inlet area which increases mass flow rate and reduces the additive drag at the lower flight Mach numbers, improving the net thrust.

Studies on variable area inlets are generally based upon translation of the spike with the cowl in fixed position. While this improves performance in level flight at zero angle of attack, it offers no improvement for overcoming thrust losses due to asymmetric flow at higher angles of attack. In this attitude, the oblique shock pattern is altered, the shock extending forward of the cowl lip on the lee side, and entering the cowl lip on the windward side; the supersonic flow is then required to make a sharp turn at the inlet, resulting in losses in total pressure recovery,

2 causing a loss in thrust when maximum thrust is needed for climb.

A novel flexible supersonic inlet is shown in FIGURE 18 which provides both a variable area inlet and .a rotatable inlet, in which the diffuser operates with maximum pressure recovery independent of angle of attack, and over a range of flight Mach numbers. The variable area is obtained by extension or retraction of a movable portion of the cowl with the spike remaining fixed; rotation is accomplished by rotating both the spike and the movable cowl in fixed relationship. The two types of motion (translation and rotation) are mutually exclusive.

In FIGURE 18, the movable cowl 381 and aft outer body 382 are joined by a flexible bellows assembly, 383, which permits freedom of motion of the movable cowl 381 in either translation or rotation. The spike assembly 384 is pivotally mounted at 385 to the aft inner body 386. A worm-gear 387 is mounted on the pivot axis 385 and driven by the actuator 3S8. Diametrically opposed struts 389 are rigidly attached to the movable cowl 381 and are slidably engaged with the parallel surfaces 390 of the spike assembly. An arm 391 passing through suitable openings 392 in the surfaces 39% is fixed to each of the opposed struts 389. A linear actuator 393 centrally mounted by suitable brackets 394 to the spike assembly 384 is attached by the pin 395 to the arm 391.

Variable area is then provided by the linear actuator 393 which extends or retracts the movable cowl 381 through the arm 391 and struts 389, the latter sliding on the parallel surfaces 390 as the inlet area is increased or decreased. This provides optimum net thrust over a range of mach numbers corresponding to the minimum and maximum forward or aft axial displacement of the cowl with respect to the spike. The internal recovery pressure acting forward on the cowl internally is somewhat balanced by the opposing pressure forces due to drag acting on the cowl externally, minimizing the magnitude of control forces required.

For flight at higher angles of attack, the entire inlet may be rotated through an angle equal and opposite to the angle of attack, the inlet remaining oriented in the direction of the relative wind. Rotation of the spike assembly by the actuator 388 through the worm-gear 387 causes the cowl 381 to rotate with it by engagement of the struts 389 with the parallel surfaces 390, the cowl and spike assembly rotating as a single rigid body about the pivot 385. Either actuator 388 or 393 may be operated independently of the other, or simultaneously. The inlet area may then be varied by the actuator 393 while in any angle of attack position, and vice versa. The oblique shock pattern then remains symmetrical about the inlet for all angles of attack, the air flow being turned internally within the duct at subsonic speeds.

The rotating inlet may, of course, be provided without the variable area inlet, and conversely, the variable area portion may be used without rotation. Also, if desired, the same principle may be applied to provide rotation of the inlet in a horizontal plane for improved inlet conditions at angle of yaw. For combined angle of attack and yaw rotation, the pivot 385 may 'be replaced by a universal joint, or gimbal ring, with two actuators lo cated 90 apart controlling rotation in each plane (not shown).

The bellows assembly 383 is shown in greater detail in FIGURE 19. In order to provide smooth exterior and interior surfaces, the convolutions may be filled with a foamed elastomeric material 397 such as silicone or poly-urethane, bonded to the metal bellows 396. The portion near the radius of each convolution as at 398 may be left blank. The skin surface 399' may be sealed in molding or inner and outer sleeves 400 of silicone or Flexible Supersonic Exits While any of the variable geometry nozzles previously described may be used with air-breathing engines, such as FIGURE 9, the flexible inlet FIGURE 36 may be adapted to operate as a flexible nozzle as well. It offers another form of variable area nozzle with vector control, particularly when used with engines operating at relatively low chamber pressures when balancing of pressure forces acting on the cowl such as previously described are not important. Proper provision for cooling would be necessary for steady state operation over long periods.

The individual actuators for the nozzle may be used as described for the diffuser, controlled by the various means previously described, with any input signal as the control parameter. Automatic nozzle variable area control, responsive'td internal fluid pressure acting on the cowl, may be provided by replacing the linear actuator 393 with a mechanical or fluid spring mounted in the same position as the actuator as described for previous nozzle types. Also, when used as an exit nozzle, the movable cowl may provide some internal supersonic expansion by suitable design similar to that described for FIGURE 9.

Improved Performance of Aircraft and Missiles With Flexible Air-Breathing Engines Improved acceleration, rate of climb and maneuverability in the rarified atmospheres of high altitude flight may be accomplished by aircraft and/ or missiles powered by improved air-breathing engines incorporating flexible inlets and exits.

Improved stability and control may be provided during take-off and landing particularly for vertical take-off and landing (VTOL), and short take-off and landing (STOL) aircraft.

The operation in flight of an aircraft and/or missile powered with the improved air-breathing engine are illustrated in FIGURES 20 and 21, in which 410 is the airframe, 411, the wing or lifting surface, and 412 are airbreathing engines. The latter may be turbo-jet, ram-jet or ducted rocket engines, using nuclear or chemical fuels and in the case of the latter may be either liquid fuels or solid propellant fuels.

The airframe houses a guidance system, control system, armament and/ or warhead, and in the case of manned aircraft, the crew. In piloted aircraft, the conventional control system may now operate the thrust vector control for the exit nozzle in place of the aero dynamic control surfaces. Pitch and yaw moments may be obtained by uni-directional rotation of the exit nozzles in the two engines, mounted outboard on the wing, While rolling moments may be provided by contra-directional rotation of the pitch control for each nozzle. In pilotless aircraft (missiles), the same system is operated by the guidance and control system.

For both aircraft and/or missiles, the flexible inlet may be controlled automatically; rotation of the inlet may be accomplished by a feed-back system utilizing an angle of attack indicator to operate the actuator 388 of FIGURE 18. The variable area inlet may be operated automatically by means responsive to pressure probes which operate the actuator 393 of FIGURE 18 so as to position the cowl in relation to the oblique shock waves corresponding to the flight mach number.

FIGURE 21 illustrates the aircraft and/or missile in a climb at an angle of attack, at. The flexible inlet of the air-breathing engine is rotated through the angle oc, thereby maintaining its heading directly into the relative wind. The exit nozzle is canted to provide a component of the thrust vector in a plane normal to the longitudinal axis to maintain the desired attitude.

The safety, reliability and range of air-launched missiles, such as long range air-to-surfa-ce missiles launched from supersonic bombers; air-to-air missiles; decoys; and

22 drones, may be improved when the flexibility of the airbreathing engine as described above is combined with the reliability of the solid propellant engine.

The ducted solid propellant rocket engine, or its equivalent, the solid fueled ram-jet engine increases reliability because of improved ignition and combustion. The solid propellant rocket, which employs a fuel rich propellant for this application becomes a hot gas generator whose fuel rich products of combustion then ignite with the oxygen of the air to complete the combustion process in the ram-jet combustion chamber. N o pressurizing means are required for fuel injection as with liquid fuels and no flame holders are required; both reactants, the incoming air-stream as well as the gaseous fuel being at elevated temperatures, increasing velocity of flame propagation and improving combustion at the low recovery pressures in high altitude flight.

A variable area nozzle may be used for the solid propellant gas generator as previously described; in this case, its purpose being to meter the flow of gaseous fuel to provide the desired mixture ratio in the ram-jet combus tion chamber under various flight conditions.

The specific impulse of the solid propellant, when augmented with free oxygen from the atmosphere is several times greater than the solid propellant which carries all of its own oxygen, thereby increasing range, or conversely, reducing the size, and weight of the missile, permitting a larger number of missiles for a given allowable weight.

The flexible inlets and/ or exits, including thrust vector control, permit improved maneuverability while eliminating large aerodynamic surfaces, reducing drag, weight, and improving storeability in flight.

Further Improvement In Nozzle Configuration Supersonic nozzles for jet engines operating at upper altitudes and in space flight have available extremely high pressure ratios even with low chamber pressure, since the ambient discharge pressure approaches zero. Extremely large area ratio nozzles are desired to provide optimum engine efliciency.

A conventional converging-diverging nozzle becomes exceedingly long and heavy as the area expansion ratio increases. While some improvement is possible by con toured nozzles having a bell-shaped expansion, these offer no improvement with respect to off-design performance when operating at lower pressure ratios, with over-expansion resulting in substantial thrust losses.

While the new plug nozzle, described in FIGURES 1 through 10, offer improvement in this respect, as the expansion ratio increases the lip diameter also increases, as represented by the line 32 of FIGURE 2. The larger diameter increases the annular throat area distributed along the larger circumference, necessitating further decrease in the normal distance from the lip to the plug surface to maintain the desired throat area. This reduced clearance between the cowl and plug increases the heat transfer rate at the throat, and makes manufacturing tolerances as well as control in flight more critical. Some improvement is possible by utilizing some internal expan- SlOIl.

A new approach is illustrated in FIGURE 22 in which large area expansion ratios may be obtained while the length and weight of the nozzle is substantially reduced; clearances at the annular throat area may be increased; and the advantage of improved off-design performance at lower pressure ratios may be retained.

These features are obtained in FIGURE 22 by applying the principles of Prandtl-Meyer flow in a reverse manher from that described for external expansion with the plug nozzle previously described. In the latter, the working fluid is directed radially inwardly towards the nozzle axis, with external expansion occuring as the fluid turned the corner at the cowl lip. In FIGURE 22, the working fluid is directed radially outwardly away from 23 the nozzle axis with internal expansion occuring as the fluid turns the corner on an internal plug.

This may be described in greater detail by referring to FIGURE 22 in which 410 may be the thrust chamber of a liquid propellant rocket engine having a fuel and oxidizer inlets at 411' and 412. An internal plug 413 supported by a rod 414 is positioned relative to the contoured expansion cowl, 415, extending from the thrust chamber 410'.

The working fluid from the thrust chamber passes radially outward through the annular channel 416 at some angle, 0, with respect to the nozzle axis, the fluid then turning through some angle, 1 the Prandtl-Meyer angle corresponding to the existing pressure ratio, about the lip 417 of the plug 413. In this design, the cowl 415 corresponds to the isentropic surface, 21, of the external expansion plug nozzle of FIGURE 2, and may be designed to conform to a streamline of fluid flow by the method of characteristics as in the case of the isentropic surface 21.

The plug lip 417 now corresponds to the cowl lip 25 of FIGURE 2 (and other figures). For lower off-design pressure ratios the flow will turn through smaller Prandtl- Meyer angles, 1 resulting in an annular jet having an internal free-stream boundary as shown by the line 418 of FIGURE 22.

In this case, the working fluid will expand only to the local ambient pressure with no over-expansion, in contrast with the conventional converging diverging nozzle where the nozzle will flow full with over-expansion reducing the pressure over the entire nozzle exit area below ambient atmospheric pressure.

In the annular jet shown in FIGURE 22, base drag will occur over the central core where no flow of the working fluid occurs. This area is separated from the ambient atmosphere by the annular jet. Due, however, to the large velocity gradient at the internal jet boundary 418, and due to the viscosity of the gases, an internal circulation will occur in which atmospheric gases will flow inwardly or forward along the nozzle axis and adjacent thereto, and outwardly, or aft, near the jet boundary as illustrated by lines 419 of FIGURE 22, thereby contributing to reduced base drag over the central core area. Also, this area is only a small portion of the total nozzle area, the remainder of which is at ambient pressure due to complete expansion; even low sub-atmospheric pressures on the central core area would produce only negligible base drag due to the low absolute pressures corresponding to the upper altitudes.

As the altitude increases, the pressure ratio and Prandtl- Meyer angle increase; as the working fluid turns through larger angles, the internal jet boundary approaches the nozzle axis until, as a limiting case, the nozzle again flows full. For this condition, the plug 413 may have appended to it a conical or isentropic surface as shown by the dotted lines at 429.

For the vacuum conditions corresponding to space flight, the pressure ratio approaches infinity and the Prandtl-Meyer angles approaches its limiting value, which may vary from 130 to 180 dependent upon the value of 'y, the ratio of specific heat for the working fluid.

The radial flow passage 4116 may be constructed to have any desired value of the angle 0, the geometry being determined for individual applications; this angle as shown in FIGURE 22 may be greater than 90, having an initial forward component thereby increasing the maximum turning angle, 11, for the particular nozzle, the cowl 415 being designed accordingly to permit turning the fluid through the larger angle.

' The thrust chamber 410', cowl 415, plug 413 and rod 414 may be regeneratively cooled or insulated as determined by detail design requirements.

' The nozzle may be readily adapted to solid propellant rocket engines by supporting the rod 414 to the case, or case extension, as represented by 410'.

Variable area may be provided by relative axial displacement of the plug or cowl, with vector control accomplished by relative rotation, employing flexible con nections as previously described.

While the description and illustrations heretofore relate to axisymmetric nozzles, the same principles may be applied to nozzles having other configurations.

It is understood that while the main elements of this invention have been described, various detail modifications may be made, such as: the use of inserts of various refractory materials in the throat sections of both the cowl and the plug; use of coatings of various thermal insulating materials within the cowl and on the case adjacent to the ports, and as liners for both the case and the plug; and other modifications such as the use of regenerative cooling, film cooling, and the like, without departing from the scope of this invention.

What is claimed is:

1. A fluid conduit thrust-producing device having a variable throat area, comprising a hollow body adapted to be supplied with an elastic working fluid under pressure, said body having a tapered end portion; a cowl surrounding and concentric with said body and having a peripheral li-p forming an annular orifice with said body adjacent said end portion; means for moving said cowl axially fore and aft said body to vary the area of said orifice for thrust control; and, means for tilting the cowl about an axis normal to the fore and aft axis of the body to elfect thrust vector control, said cowl forming a fluidconductive passage with said body connecting the interior of said body with said annular orifice, and said cowl having a member separable therefrom and defining a portion of said passage, for effecting thrust termination.

2. A fluid conduit thrustaproducing device in accord ance with claim 1 including flexible sealing means flexibly sealingly connecting said cowl to said body.

3. A fluid conduit thrust-producing device in accordance with claim 2 in which said flexible sealing means comprises an annular bellows member encasing said body portion and sealingly connected at one end to said body portion and at the other to said cowl.

4. A fluid conduit thrust-producing device in accordance with claim 2 in which said flexible sealing means comprises an annular bellows member and annular restraining members interposed between the folds of the bellows member.

5. A fluid conduit thmst-producing device in accordance with claim 2 in which said flexible sealing means comprises a double-ended folded sleeve having spacedapart wall portions joined together at one pair of corresponding ends by a loop portion and sealingly connected at the. other pair of corresponding ends to said cowl and said body portion, respectively.

6. A fluid conduit thrust-producing device in accordance with claim 1 in which said separable member is electrically actuated.

7. An airborne vehicle comprising an air-frame, at least one air-breathing jet engine carried by the air-frame, a variable geometry plug nozzle for said engine having thrust vector control for maneuverability and stability of the vehicle, said variable geometry plug nozzle comprising an axially-disposed plug, a cowl surrounding said plug and concentric therewith, the cow-l being spaced from the plug to define a throat inclined to the plug axis, through which gases from the engine flow aft between the cowl and the plug, and the cowl also being mounted for angular movement on each of a pair of thrust vector control axes which are normal to each other and also normal to the plug axis, and means for moving said cowl on said vector control axes to obtain such thrust vector control.

8. An airborne vehicle in accordance with claim 7 said cowl being mounted 'for axial reciprocation relative to the nozzle plug to provide a variable throat area by reciprocal movement of the cowl with respect to the nozzle plug.

9. An airborne vehicle comprising an air-frame, at least one air-breathing jet engine carried by the air-frame, a variable geometry inlet diffuser for said engine, said diffuser comprising a diffuser spike, an air-inlet cowl surrounding said spike and concentric therewith, the cowl being spaced from the spike to define a throat inclined to the spike axis, through which air flows aft to the engine between the cowl and the spike, the cowl being mounted to the spike for fore and aft translational movement relative to the spike, and the spike and cowl being mounted for angular movement as a unit on a rotational axis normal to the spike axis, means for effecting said translational movement of the cowl, to vary the throat area, and means for effecting said angular movement of the cowl and spike to vary the inclination of the throat and the direction of the spike axis; and, a variable geometry exit nozzle for said engine, said exit nozzle comprising an axially disposed plug, a cowl surrounding said plug and concentric therewith, the cowl being spaced from the plug to define a throat inclined to the plug axis, through which gases from the engine flow aft between the cowl and the plug, the cowl also being mounted for angular movement on each of a pair of thrust vector control axes which are normal to each other and also tothe plug axis, to obtain such thrust vector control.

10. A missile having a sustainer rocket and a disposable solid propellant booster rocket separably connected to the sustainer rocket, said booster rocket comprising a propellant case, a variable throat area nozzle; timecontrollcd means for increasing the nozzle throat area after a selected time interval to reduce case pressure and terminate thrust; means for separating the booster rocket from the sustainer rocket and, time-controlled means energized by separation of the booster rocket from the sustainer rocket, for decreasing the nozzle throat area after a selected time interval so as to increase the case pressure to a bursting pressure.

11. A jet nozzle for a reaction combustion engine, including, in combination: means providing a chamber for propellant gases; an axially disposed nozzle plug for directing the flow of propellant gases through the aft end of the engine; a cowl surrounding the plug and concentric therewith, the cowl being spaced from the plug to define a throat transverse to the plug axis, through which the propellant gases flow aft between the cowl and the plug; means for directing propellant gas pressures against the cowl in both forward and aft directions to reduce the force required to move the cowl axially against gas pressures; and, means for moving the cowl axially against the resultant gas pressures acting thereon.

12. In the nozzle of claim 11, in combination, means for diverting the fiow of gas from the aft direction to a transverse direction to terminate nozzle thrust.

13. In the nozzle of claim 12, said diverting means comprising normally-closed ports for lateral escape of gas from the cowl, and means for opening said ports.

14. In a reaction combustion engine, in combination: means providing a chamber for propellant gases; orificeforming members providing an annular orifice at the aft end of said chamber for expelling said gases, one of the members forming a perimeter of the orifice and being fixed, and another of said members forming the other perimeter of said orifice and being axially movable with respect to said fixed member to vary the area of the orifice; means for applying propellant gas pressures to said movable member in the forward and aft directions to reduce the force required to move said member axially; and, means for applying a minor axial force to move said member axially against the resultant gas pressure acting thereon to vary the area of the orifice.

15. In a reaction combustion engine, in combination, shell means providing a chamber for propellant gases under pressure; a nozzle comprising an annular propellant gas escape orifice formed by a central plug member and a member constituting part of the surrounding shell of said chamber, one of said members having a part which is axially-movable while the other member is fixed, to vary the area of said orifice; and, means for applying forward and aft propellant gas pressures to a portion of the surface of one of said orifice-forming members to reduce the force required for axial movement of said movable part. i 16. In the combination of claim 15, means for imparting axial control movements to said movable part, including an actuating rod extending longitudinally along the chamber, a connection between said rod and said axiallymovable part, and means external to the chamber for controllably moving said actuating rod.

17. In a reaction combustion engine, in combination, means providing a chamber for propellant gases; a nozzle at the aft end of said chamber for controllably expelling said gases; said nozzle comprising 'an annular throat orifice formed between a centrally-disposed, stationary, tapered plug and an axially movable, circumferential cowl at the aft end of the chamber surrounding and spaced from said plug; and, means for varying the throat area of said orifice by controlled axial movements of the cowl with respect to the stationary plug, said throat varying means including means for directing propellant gas pressures against the surface of the cowl in both forward and aft directions to provide a reduced resultant gas pressure opposing axial movement of said cowl, and means for controllably axially moving said cowl against said resultant pressure.

18. The combination of claim 17, wherein the means for directing said forward and aft gas pressures against the cowl include gas ports in the chamber wall forward of said throat orifice and communicating with the interior of said cowl.

19. The combination of claim 18 wherein the interior of said cowl is hollowed to provide surfaces reacting to forward and aft gas pressures, respectively.

20. A variable area jet nozzle including, in combination: means forming a chamber for propellant gases; a plug having an isentropic surface; a cowl surrounding and concentric with the plug and spaced therefrom to define a throat transverse to the plug axis, through which propellant gases from said chamber flow, the cowl being axially movable with respect to the plug, means for leading propellant gases from the interior of the chamber to exert pressure on the cowl in both forward and aft directions to reduce the force required to move it axially against gas pressures and means for moving the cowl axially against the resultant gas pressures acting thereon.

21. In a reaction motor for flying objects, in combination: means forming a chamber for propellant gases; a plug centrally disposed in and projecting from the aft end of said chamber; a cowl at the aft end of said chamber having a lip surrounding and spaced from the plug to form a throat orifice for escape of propellant gases from said chamber, the axis of the lip of the cowl being normally coincident with that of the plug to cause the escaping gas to issue in a stream parallel to the axis of the plug; and, means for tilting the cowl to dispose its lip axis at an angle to that of the plug to cause the stream of propellant gases to issue at an angle to the plug axis and thereby change the direction of flight of the object propelled by said motor.

22. The combination of claim 21, the cowl being movable axially forward and aft to control the area of said orifice.

23. The combination of claim 22, including control means for imparting axial movement to said cowl.

24. A reaction motor for flying objects, including, in combination: means forming a propellant gas chamber; a plug at the aft end thereof; a circular cowl surrounding and concentric with said plug to form a nozzle having an annular orifice; and, means for tilting the cowl about an axis normal to the fore and aft axis of the plug to effect vector control of the jet stream issuing from the nozzle.

27 28 25. The motor of claim 24, said cowl also being mov- 2,583,570 Hickman Ian. 29, 1952 able axially forward and aft to control the area of said 2,603,433 Nosker July 15, 1952 orifice. 2,683,349 Lawrence July 13, 1954 26. In the motor of claim 24, the combination of 2,683,962 Grifiith July 20, 1954 means for applying forward and aft gas pressures .to the 2,686,473 Vogel Aug. 17, 1954 cowl to facilitate ease of movement thereof. 2,701,441 Mitchell Feb. 8, 1955 27. The motor of claim 25, including, the combination 2,724,947 Meyer Nov. 29, 1955 of means for applying forward and aft gas pressures to 2,746,242 Reed May 22, 1956 the cowl to facilitate ease of movement thereof. 2,746,243 Pitt et a1. May 22, 1956 2,760,336 Reniger Aug. 28, 1956 References Cited in the file of this patent 6 ,152 lss hugzl t a1 Sept. 11, 1956 2,776,8 ren a Jan. 8, 1957 UNITED STATES PATENTS 2,780,914 Ring Feb. 12, 1957 254,048 Robertshaw Feb, 21, 1882 2,789,505 Cumming et p 3, 1957 2 4 317 M Ti h et 1, S 12 1 2 2,810,533 Lauderdale et a1. Oct. 22, 1957 2,406,560 Pope Aug. 27, 1946 2,811,827 Kress 1957 2,413,621 Hammond Dec. 31, 1946, 2,826,895 English 1958 2,418,488 Thompson Apr, 8, 1947 2,841,953 T998119 y 1958 2,478,958 Wheeler Aug. 16, 1949 2,841,957 Thorpe y 8, 1958 ,5 7 Chandler Man 7, 1 5 2,865,169 Hausmann 2 1 2 503 310 Weiss A 11 1950 2,928,235 JOhIISOIl M l 15, 196 2 505 79 ki May 2 1950 2,932,945 Brandt P 1960 2,524,591 Chandler Oct. 3, 1950 FOREIGN PATENTS 215525497 f et a1 May 1951 5,099 Great Britain Dec. 12, 1878 2,570,629 AnXwImaZ 91 1951 2. 757,457 Great Britain Sept, 19, 1 5 2,571,386 Sarnoff Oct. 16, 19 1 1,003,758 France Nov. 21, 1951 2,578,202 Palme Dec. 11, 1951 1,098,274 France Mar. 2, 1955

Claims (1)

11. A JET NOZZLE FOR A REACTION COMBUSTION ENGINE, INCLUDING, IN COMBINATION: MEANS PROVIDING A CHAMBER FOR PROPELLANT GASES; AN AXIALLY DISPOSED NOZZLE PLUG FOR DIRECTING THE FLOW OF PROPELLANT GASES THROUGH THE AFT END OF THE ENGINE; A COWL SURROUNDING THE PLUG AND CONCENTRIC THEREWITH, THE COWL BEING SPACED FROM THE PLUG TO DEFINE A THROAT TRANSVERSE TO THE PLUG AXIS, THROUGH WHICH THE PROPELLANT GASES FLOW AFT BETWEEN THE COWL AND THE PLUG; MEANS FOR DIRECTING PROPELLANT GAS PRESSURES AGAINST
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Cited By (34)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
US3210937A (en) * 1962-04-10 1965-10-12 Jr Henry A Perry Thrust control apparatus
US3214906A (en) * 1962-07-05 1965-11-02 Aerojet General Co Hybrid rocket motor
US3218798A (en) * 1963-01-30 1965-11-23 Atlantic Res Corp Spherical booster
US3234731A (en) * 1962-01-10 1966-02-15 North American Aviation Inc Variable thrust device and injector
US3279183A (en) * 1962-09-07 1966-10-18 United Aircraft Corp Vectorable plug cluster nozzle rocket
US3286462A (en) * 1963-10-09 1966-11-22 Thiokol Chemical Corp Gas generator having slow burning grain for variable gas flow
US3286474A (en) * 1962-12-05 1966-11-22 North American Aviation Inc Hoop segmented injector and combustor
US3289946A (en) * 1963-08-07 1966-12-06 Gen Electric Annular convergent-divergent exhaust nozzle
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US3314609A (en) * 1962-09-07 1967-04-18 United Aircraft Corp Vectorable plug cluster nozzle rocket
US3353357A (en) * 1965-01-18 1967-11-21 Thiokol Chemical Corp Rocket powerplant
US3446436A (en) * 1966-11-29 1969-05-27 Thiokol Chemical Corp Rocket thrust nozzle system
US3447325A (en) * 1966-08-15 1969-06-03 Bristol Siddeley Engines Ltd Controlling supersonic air intakes
US3486698A (en) * 1966-07-13 1969-12-30 Thiokol Chemical Corp Roll and directional control apparatus for rocket motors
US3489373A (en) * 1967-01-03 1970-01-13 Arthur R Parilla Missile configurations,controls and utilization techniques
US3760589A (en) * 1969-03-24 1973-09-25 Thiokol Chemical Corp Throttling mechanism for controlling the thrust level of a solid propellant rocket motor
US3806064A (en) * 1968-10-03 1974-04-23 A Parilla Missile configurations, controls and utilization techniques
US4363445A (en) * 1979-11-23 1982-12-14 Societe Nationale D'etude Et De Construction De Moteurs D'aviation, "S.N.E.C.M.A." Thrust-vectoring nozzle for jet propulsion system
FR2549146A1 (en) * 1983-07-11 1985-01-18 Europ Propulsion Propulsion to ramjet missile propellant acceleration INTEGRATED
US5125596A (en) * 1989-05-23 1992-06-30 Cavalleri Robert J Fluid shielded movable strut for missile and rocket thrust vector control
EP0554771A1 (en) * 1992-01-29 1993-08-11 Thiokol Corporation Method for the use of controlled burn rate, reduced smoke, solid propellant formulations
EP0646716A1 (en) * 1993-09-30 1995-04-05 Societe Europeenne De Propulsion Bidimensional vectoriable nozzle
FR2780449A1 (en) * 1998-06-29 1999-12-31 Snecma Missile/satellite vectored thrust device
EP1735531A2 (en) * 2003-12-01 2006-12-27 The University Of Mississippi Method and device for reducing engine noise
EP1813907A1 (en) * 2006-01-26 2007-08-01 Deutsche Forschungsanstalt für Luft- und Raumfahrt e.V. Missile for the supersonic range
EP1931871A2 (en) * 2005-09-13 2008-06-18 Aerojet-General Corporation Thrust augmentation in plug nozzles and expansion-deflection nozzles
US20090084888A1 (en) * 2005-03-29 2009-04-02 Mordechai Shai Steering system and method for a guided flying apparatus
US20100162684A1 (en) * 2008-12-26 2010-07-01 Von David Baker Aircraft nozzle
RU2527228C1 (en) * 2013-02-19 2014-08-27 Федеральное государственное бюджетное образовательное учреждение высшего профессионального образования "Тульский государственный университет" (ТулГУ) Solid-propellant engine nozzle unit
US20140331682A1 (en) * 2012-11-08 2014-11-13 Mark Bovankovich High-speed-launch ramjet booster
US20160177873A1 (en) * 2013-07-13 2016-06-23 Mbda Uk Limited A thrust flow powered vehicle
US9810178B2 (en) 2015-08-05 2017-11-07 General Electric Company Exhaust nozzle with non-coplanar and/or non-axisymmetric shape
FR3068737A1 (en) * 2017-07-07 2019-01-11 Arianegroup Sas Configuring propeller for delivering modular thrust
RU2681733C1 (en) * 2017-12-28 2019-03-12 Акционерное общество "Конструкторское бюро химавтоматики" Camera lpr

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Cited By (45)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
US3312068A (en) * 1960-12-05 1967-04-04 North American Aviation Inc Horizontal flow thrust chamber
US3234731A (en) * 1962-01-10 1966-02-15 North American Aviation Inc Variable thrust device and injector
US3210937A (en) * 1962-04-10 1965-10-12 Jr Henry A Perry Thrust control apparatus
US3214906A (en) * 1962-07-05 1965-11-02 Aerojet General Co Hybrid rocket motor
US3279183A (en) * 1962-09-07 1966-10-18 United Aircraft Corp Vectorable plug cluster nozzle rocket
US3314609A (en) * 1962-09-07 1967-04-18 United Aircraft Corp Vectorable plug cluster nozzle rocket
US3286474A (en) * 1962-12-05 1966-11-22 North American Aviation Inc Hoop segmented injector and combustor
US3218798A (en) * 1963-01-30 1965-11-23 Atlantic Res Corp Spherical booster
US3289946A (en) * 1963-08-07 1966-12-06 Gen Electric Annular convergent-divergent exhaust nozzle
US3286462A (en) * 1963-10-09 1966-11-22 Thiokol Chemical Corp Gas generator having slow burning grain for variable gas flow
US3353357A (en) * 1965-01-18 1967-11-21 Thiokol Chemical Corp Rocket powerplant
US3486698A (en) * 1966-07-13 1969-12-30 Thiokol Chemical Corp Roll and directional control apparatus for rocket motors
US3447325A (en) * 1966-08-15 1969-06-03 Bristol Siddeley Engines Ltd Controlling supersonic air intakes
US3446436A (en) * 1966-11-29 1969-05-27 Thiokol Chemical Corp Rocket thrust nozzle system
US3489373A (en) * 1967-01-03 1970-01-13 Arthur R Parilla Missile configurations,controls and utilization techniques
US3806064A (en) * 1968-10-03 1974-04-23 A Parilla Missile configurations, controls and utilization techniques
US3760589A (en) * 1969-03-24 1973-09-25 Thiokol Chemical Corp Throttling mechanism for controlling the thrust level of a solid propellant rocket motor
US4363445A (en) * 1979-11-23 1982-12-14 Societe Nationale D'etude Et De Construction De Moteurs D'aviation, "S.N.E.C.M.A." Thrust-vectoring nozzle for jet propulsion system
FR2549146A1 (en) * 1983-07-11 1985-01-18 Europ Propulsion Propulsion to ramjet missile propellant acceleration INTEGRATED
US4631916A (en) * 1983-07-11 1986-12-30 Societe Europeenne De Propulsion Integral booster/ramjet drive
US5125596A (en) * 1989-05-23 1992-06-30 Cavalleri Robert J Fluid shielded movable strut for missile and rocket thrust vector control
US5579634A (en) * 1992-01-29 1996-12-03 Thiokol Corporation Use of controlled burn rate, reduced smoke, biplateau solid propellant formulations
EP0554771A1 (en) * 1992-01-29 1993-08-11 Thiokol Corporation Method for the use of controlled burn rate, reduced smoke, solid propellant formulations
FR2710694A1 (en) * 1993-09-30 1995-04-07 Europ Propulsion Nozzle symmetrical two-dimensional planar expansion-deflection adjustable thrust and application in a spacecraft.
EP0646716A1 (en) * 1993-09-30 1995-04-05 Societe Europeenne De Propulsion Bidimensional vectoriable nozzle
WO2000000731A1 (en) * 1998-06-29 2000-01-06 Societe Nationale D'etude Et De Construction De Moteurs D'aviation -S.N.E.C.M.A.- Compact and adjustable tailpipe for piloting aerospace craft
US6543717B1 (en) 1998-06-29 2003-04-08 Societe Nationale D'etude Et De Construction De Moteurs D'aviation S.N.E.C.M.A. Compact optimal and modulatable thrust device for controlling aerospace vehicles
FR2780449A1 (en) * 1998-06-29 1999-12-31 Snecma Missile/satellite vectored thrust device
EP1735531A2 (en) * 2003-12-01 2006-12-27 The University Of Mississippi Method and device for reducing engine noise
EP1735531A4 (en) * 2003-12-01 2015-01-21 Univ Mississippi Method and device for reducing engine noise
US20090084888A1 (en) * 2005-03-29 2009-04-02 Mordechai Shai Steering system and method for a guided flying apparatus
US8080771B2 (en) * 2005-03-29 2011-12-20 Israel Aerospace Industries Ltd. Steering system and method for a guided flying apparatus
EP1931871A2 (en) * 2005-09-13 2008-06-18 Aerojet-General Corporation Thrust augmentation in plug nozzles and expansion-deflection nozzles
EP1931871A4 (en) * 2005-09-13 2013-04-17 Aerojet General Co Thrust augmentation in plug nozzles and expansion-deflection nozzles
EP1813907A1 (en) * 2006-01-26 2007-08-01 Deutsche Forschungsanstalt für Luft- und Raumfahrt e.V. Missile for the supersonic range
US20070295856A1 (en) * 2006-01-26 2007-12-27 Deutsches Zentrum Fur Luft-Und Raumfahrt E.V. Flying object for transonic or supersonic velocities
US7775480B2 (en) 2006-01-26 2010-08-17 Deutsches Zentrum Fur Luft-Und Raumfahrt E.V. Flying object for transonic or supersonic velocities
US20100162684A1 (en) * 2008-12-26 2010-07-01 Von David Baker Aircraft nozzle
US8459036B2 (en) * 2008-12-26 2013-06-11 Rolls-Royce Corporation Aircraft nozzle having actuators capable of changing a flow area of the aircraft nozzle
US20140331682A1 (en) * 2012-11-08 2014-11-13 Mark Bovankovich High-speed-launch ramjet booster
RU2527228C1 (en) * 2013-02-19 2014-08-27 Федеральное государственное бюджетное образовательное учреждение высшего профессионального образования "Тульский государственный университет" (ТулГУ) Solid-propellant engine nozzle unit
US20160177873A1 (en) * 2013-07-13 2016-06-23 Mbda Uk Limited A thrust flow powered vehicle
US9810178B2 (en) 2015-08-05 2017-11-07 General Electric Company Exhaust nozzle with non-coplanar and/or non-axisymmetric shape
FR3068737A1 (en) * 2017-07-07 2019-01-11 Arianegroup Sas Configuring propeller for delivering modular thrust
RU2681733C1 (en) * 2017-12-28 2019-03-12 Акционерное общество "Конструкторское бюро химавтоматики" Camera lpr

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