Патент USA US3035806код для вставки
May 22, 1962 c. A. GLASS 3,035,796 DUAL THRUST ROCKET BOOSTER TUBE Filed NOV. 2l, 1958 2 Sheets-Sheet 1 / wmo BLAST 26 23 w CECIL INVENTOR. A. GLASS BY A 4. /f ATTORN Ys. May 22, 1962 c. A. GLASS 3,035,796 DUAL THRUST ROCKET BOOSTER TUBE Filed NOV. 2l, 1958 2 Sheets-Sheet 2 l 36 d0 aV d 28 ‘Il1|' m 42 46 INVENTOR. CECIL BY ATTORNEYS. A. GLASS United States Patent 0f ice Patented May 22, 1962 1 2 3,035,796 The largest volume of sustainer propellant can be used by employing a cylindrical internal burning propellant grain having a central internal perforation extending be DUAL THRUS'I‘ ROCKET BOOSTER TUBE Cecil A. Glass, China Lake, Calif., assigner to the United tween its ends which is bonded to or rests snugly against States of America as represented by the Secretary of the Navy t 3,035,796 the motor tube on its outer (non-burning) surface. This propellant extends the full length of the rocket motor tube. With the tandem arrangement of this device, the “tube action” propellant takes up a substantial portion of the available motor tube length, leaving only a portion of Filed Nov. 21, 1958, Ser. No. 775,662 4 Claims. (Cl. 244-122) (Granted under Title 35, U.S. Code (1952), sec. 266) rIhe invention described herein may be manufactured 10 length of the envelope for the sustainer grain, thereby preventing attainment of maximum power. Furthermore, and used by or for the Government of the United States in the present‘state of the art, this particular margin of of America for governmental purposes without the pay power is very critical in attaining reliable ground level ment of any-royalties thereon or therefor. ejection capacity. Also, the separate microswitch circuit This invention relates to combined inner tube action for igniting the :free ñight propellant grain increases the possibility of malfunction or involuntary firing. and rocket action aircraft ejection seats and in particular to those ejection seats which incorporate the inner tube and rocket motor into a dual thrust single assembly. A short discussion of free volume within the internal perforation of a cylindrical hollow grain will give a better understanding of the present invention. The free volume . FIG. l generally illustrates a typical combined tube action and rocket action ejection seat. It is used princi pally in connection with military aircraft. The term 20 or void space within an internal burning propellant pro “tube action” and “rocket action” are derived from the two types of propulsion used in ejection. The seat is propelled the few feet necessary to leave the cockpit lby tube action, that is, contained combustion gases within a launching tube Working against a sliding member or piston 25 which has a known area. In’dual thrust rocket termi nology thisís called the “booster” phase of propulsion._ After leaving the cockpit, the seat is propelled by free flight rocket action. In dual thrust rocket terminology vides the exit channel for the passage of the products of combustion in reaching the nozzle. Depending upon the particular propellant and its configurations this free vol ume must equal 0.1 to 0.4 of the total volume of the com bustion chamber for satisfactory performance. Anything affecting the use of this free volume as an exit channel, such as the presence of objects that otherwise occupy the space, causes unsatisfactory performance. The constric tion of the exit channel causes an increase in velocity of this .is called the sustainer phase of propulsion. It is the 30 the hot combustion gases resulting in excessive scrubbing and wearing of the propellant surface, which is termed goal in the design of such ejection seats to provide free » flight travel to an elevation permitting safe parachute “erosive burning” and subsequent excessive motor tube escape, even when ejection is initiated at ground level, as in take-off and landing situations. pressures. ` 'Two‘prior art devices are known and will be described. In one device the combustion gases for the lbooster or tube tion comprises a simplified economical structure which Generally stated, the improvement of the present inven permits the combination of a booster tube or inner cham ber, for producing the gases for the tube action, within the internal perforation of an elongated hollow internal burning grain, while later making effective use of the space chamber leads into a port in the breech end of the launcher tube so that the combustion ` gases expand 40 occupied by the booster tube as free volume during sus action are produced in a separate combustion chamber located externally of the launcher tube. The separate through the launcher tube. The piston and free flight rocket are combined in a single «assembly which slideably engages within the launcher tube. Ignition of the free flight rocket is effected by a separate barrier assembly tainer grain ñring. The effectiveness of the present inven tion is evident when it is realized that the typical internal burning propellant grain suitable for the above referred to envelope size has an internal perforation diameter of 11A which is slideably engaged in the launcher tube and 45 inch. It has been found possible, as a result of this inven which travels »outwardly with the piston. Initially in the tion, to place a 1 inch O_D. booster tube inside this 1% tube action, this barrier prevents the hot combustion inch internal perforation without causing unsatisfactory gases produced in the separate chamber from reaching the performance through erosion burning. rocket nozzles. However, the barrier assembly contains It is an object of the present invention to provide an a hole with a plug which is mechanically sheared olf `just 50 upward directed aircraft ejection seat with booster tube prior to the piston and rocket motor assembly leaving thelauncher tube. Hot gases pass through the hole and ¿ »and free flight rocket in the same assembly, with suffi cient power to permit reliable ground level ejection then into the rocket motor nozzle, -igniting the free flight capacity. . propellant. In this device the separate combustion cham It is another object of the present invention to provide -ber takes up critical space within the aircraft and adds 55 to the cost of manufacture. The barrier assembly is a separate moving part, which increases the possibility of malfunction. ‘ - within an aircraft ejection seat system having a given , envelope size for the booster and free flight propulsion means, more ejection power than heretofore available. It isa further object of the present invention to provide In -another prior art device the com-bustion chamber for producing gases for the tube action, the combustion cham 60 a dual thrust rocket motor wherein a booster tube may be ber for producing gases for the free flight rocket action, disposed within the internal perforation of an internal burn and the piston are combined into a single assembly. In ing propellant grain without sacrificing the use of the other words the piston is also a dual thrust rocket. The space occupied by the booster tube as effective free volume two propellant charges are disposed in tandem, that is, during sustainer grain firing. one behind the other within the length of the rocket motor. 65 A still further object of the present invention is to pro In this device, the free Hight propellant is ignited elec vide within a dual thrust rocket a >simple means, requiring trically by microswitch means, just prior to the time the rocket motor leaves the launcher. The tandem arrange no extra parts, for the timed ignition of the sustainer pro pellant. ,ment results in substantially less space for available rocket Other objects and many of the attendant advantages of motor power. It is generally recognized that the envelope 70 this invention will be readily appreciated as the saine size available for a seat ejection power system is -a space becomes better understood by reference to the following approximately 31/2 xfeet long and having a 3-inch diameter. 3,035,793 detailed description when considered in connection with wise cause undesired 4tumbling during free ñight and the -accompanying drawings wherein: secondly, yto provide a forward thrust component to soften the rearward- impact which occurs duringa high speed FIG. 1 illustrates the use of lan aircraft ejection seat with the present investiert. - . ejection when the pilot is abruptly catapulted into a high ~ FIG.. 2 is a longitudinal eentral eeetien efY the launcher tube. reeket meter tube, au@ buuster tube ,asserublnk PIG. 3 is an enlarged seetieu taken. .alena .line .3v-ê. _ef FIG. 2, ` ries. 4A, 4a and 4c are sesame 'a ius Seguential conditions during eis speed wind blast `as illustrated in FIG. 1. Nozzles 42 con tain tightly ñtted blow out plugs `44, which remain iixed in these nozzles until the sustainer propellant grain Z8 is _ Ígllîîçd.- Theseplugs are made oiga relatively heat>V resistant aantast-._ ' ' material such as aluminum and prevent premature igni ti-on of the sustainer propellant grain by entry of hot tube Y V"Ille maier Components of acera> >.ined tubeaetien .aud " gases inte Vthe main chamber through the nozzle parte reeltet aetiuu airerait eieetieu seat will. be, apparent. fr0-ul au examination of the illustrative example. 'shawn iu the? drawing,- A launcher tube 1,0y having la 1‘l1-teach .chamberY Nozzle 46lcommunic`ates with the booster tube through an 1_2 iS Secured t0 the aircraft frame, by `any suitable trunnion elbow euuueetiOu-‘Stltur exhausting the-eerrlbustiun eases of the booster tube. lA conventional nozzleclosure di aphragm 48 made et neoprene. ,is similarlyftttediuteuuz 1,1. A rocket motor Iassembly 2_0, having the, dQllble func- ` zle 46 to prevent dust or moisture from entering the booster tion of piston and thrust rocket motor, is slideably en ¿aged within the launcher tube.' The assembly includes tube. the main motor tube 2,2., a smaller inner booster tube 30 and a nozzle assembly 40. The ejection seat 23 is rigidly ' ' It is >to be understQod'that-the pilot initiates seat eiee. 20 tion by actuating some, conventional mechanism that tigres, the ignite: 32 which in turn ignites thebooster propellant attached to the motor tube 22 by means of a »suitable strips 36, 36' withinrfthe booster. tube. Typical flow bracket 2,4 and clamping ring 25. paths of the combustion gases at this moment are shownn - " Now referring to the rocket motor assembly 20 shown by the arrows in FIG. 4A. = Since theapertures 34».alo1:lg> in FIGS. 2. and 3, the Outer casing or motor Atube 22, the ' the length of the booster 30 are covered by a membrane. nose cap 2,6, and the nozzle assembly 4t2, form a main 25 consisting of the propellant strips and since as a point of design the web> thickness and strength of the propellant yis designed to have an'initial Yrupture strength greater than the booster tube combustion pressure, the gases .are C911, 28 is bonded tot or rests snugly against the insidewall of s_trained'to ilow through-the booster tube nozzle 46 intQ the motor tube 22y and is mechanically supported in the 30 the launcher tube breech chamber 12.. These gases will combustion chamber to allow for thermal expansion by work against the rocket .motor assembly 20 and provide means of resilient spacers, springs or other structural de the tube action while blow-out plugs 44. prevent these vices known in the art. ' gases from entering Vnozzles 42. As understood in the A booster tube 30 is disposed within the central per art, the cockpit canopy is cast oñ, the pilot is constrained combustion chamber for producing the sustainer or free flight combustion gases. The outer non-burning surface of a hollow cylindrical internal burning '_sustainer grain f_oration of the propellant grain 2,8, its ends being rigidly 35 to the seat, and the seat, being rigidly secured to the secured to the nose> cap Z6 and nozzle assembly 40. VThe Y rocket motor .assembly 20, YtravelsY upwardly, guided by outer diameter of tube 30 is slightly smallerY than theref rails. fective diameter of the propellant perforation so as to leave free volume within the main combustion chamber all around the booster tube, The end of the booster tube adjacent the nose cap 2.6 contains a conventional> igniterv 32 and is closed -so that the booster tube forms a small inner combustion chamber for the production of the gases for the ejection of the rocket rrrotor ‘assemblyV and Vman seat out-of the launcher. Although not shown in the draw ing, suitable igniter firing accessories such as sear' and 40 45 firing pin may be incorporated into the nose> lcap. The. booster tube has a plurality of apertures 3.4. At- the start o_f ejection these apertures are coveredby :booster fp_ro i a v . v. As the rocket motor assembly travels up the launcher tube, burning takes place normal >to the exposed burning surfaces of the propellant strips 36, ¿36C The rocket rnotor assembly 20 travels up the launcher tube until it reaches a «point where it is Within a few inches of leav ing the tube as illustrated in FIG. 4B.V At that moment the membrane over the booster tube apertures 34, consist ing of the propellant stripsa36, 36' rupture under the booster tube combustion pressure as the result of reduc tion in the web thickness of the propellantstrips 36, 36.’ by burning; The precise timing of this event is predeter mined by the- choice of initial web thicknessfor a propelf pellant strips 36„ 36', as best shown in FIGQ3, which have 50 lant s_triphaving a givenburning rateand given rupture a semi-circular section and areV bondedto the linside wall strength characteristics. The flow paths oÍf the combus 0f the beoster tube- Thusthe web. thieliuese Qi these pre tion gases at thisv moment are shown by the arrows in pellant strips form a membrane over the holesV and only FIG. 4B. -A portion/of the hot booster tube gases ñoW the inner surface of- the propellant is exposed forY burn through holes 34 into the free volume within the main ing within the booster tube. The choice of the particular 55 combustion chamber. This’ moment marks the begin booster propellant is dictated by the desired burning char acteristics. Double base formulations havinglmesa-burn ing characteristics, i.e., uniform burning rate over a rela- „ tively wide temperature range have been found to work elîectively. In the exemplaryl form, the strips are bonded 60 to the inner wall by a bonding Iagent 3:8V consisting of cot ton tape, as shown in FIG. 3, that had been dipped in polysiloxane adhesive. As is common in ftheV art, inhibit ing tape 39, bonded to the propellant, may be used to inhibit a portion of its exposed burning surface so as to 65 obtain particular burning characteristics. ' v ning of the transition between tube action Vand rocket action. It is torbe noted at this> moment, justV prior to ignition of- the sustainer the Vblow-out plugs 44 are still in place and a portion of the booster gases is still exhaust ing through'nozzle 46. . ‘ ' ' ' After a very brief ignition delay, for no more than the short time required for the piston and'rock’et motor as sembly to travel the remaining few inches within the launcher tube, the Ysustainer propellant, 28 ignites and blow-outvplugs 44 are ejected from nozzles V42.V The flow paths ¿of the sustainer propellant combustion gases an instant after ignition are s_hown in FIG.'4C,_ thebooster A multiport nozzle `40 containing several convergent divergent type nozzles 4,2A communicating Vto, the main propellant strips having burnedY away( A portion of the combustion chamber and ya similar nozzle communicat gases exhaust through n_ozzles 42,V passing'through the ing to the booster tube is secured to the end of the motor 70 free >volumes between the burning surface of the propel tube. The nozzle is canted at an angle which results in the axes of the nozzles passing near the center of gravity of the combined man-seat mass as shown in'FIG. 1'.> The purpose of this is to avoid applying unwanted unbalanced lant grain and the booster tube._; jSimultaneouslyÍ another portion ofthe combustion gases passesthrough 'the aper tures 34_into fthe boostentubeand exhaust y‘through booster tube nozzle 46„ ‘thus lusing the freevolume of torque to the man-seat combination which would other-ÍY 75 the booster tube >.during sustainer .grain tiring. This 3,035,796 6 marks the end of the transition from tube action to rocket action and the combined pilot and seat is now propelled as a free flight reaction rocket vehicle and the direction of thrust has changed from the direction of arrow T to to the piston assembly, the piston assembly being further adapted to provide combustion gases for the ejection of arrow T’ wherein the resultant passes near the center of piston assembly leaves the launcher tube; the improve gravity as shown in FIG. 1. the piston assembly out of the launcher tube and com bustion gases for reacting rocket free flight after the ' ment wherein said piston assembly comprises, an elon After the ejection seat has been shot high enough to gated main combustion chamber having a first propellant make a safe parachute escape, delay initiators release grain therein and adapted to produce said combustion the pilot from the seat and his attached parachute auto gases for reaction rocket free flight, an elongated inner matically opens by means heretofore known in the art. 10 combustion chamber having a second propellant therein Obviously many modifications and variations of the and adapted to produce the said combustion gases for present invention are possible in the light of the above the ejection of the piston assembly out of the launcher teachings. It is therefore to be understood that within tube, said inner chamber being disposed concentrically the scope of the appended claims the invention may be within the main chamber and directly communicating practiced otherwise than as specifically described. with the interior of said launcher tube to pressurize same I claim: with gases produced by said second propellant, aperture ,Y 1,. In apparatus for use in a tube action rocket action means extending through the wall of the inner chamber, aircraft ejection seat having a launcher tube attached to said aperture means adapted to be closed at the time ejec the aircraft and rocket motor assembly attached to the -tion starts, means for opening the aperture means after seat and slideably engaged within the launcher tube; the 20 the piston assembly has traveled a predetermined distance improvement wherein said rocket motor assembly com along the launcher tube thereby permitting the combus prises; an elongated rocket motor tube, an internal burn tion gases produced in the inner chamber to pass into and ing cylindrical propellant grain disposed within the ma ignite the propellant in the main combustion chamber, jority of the length of said motor tube having a central said aperture means being further adapted to permit a perforation extending between its ends, one end of said 25 portion of the gases produced in the main chamber to motor tube having a first nozzle means for exhausting pass into and exhaust through the inner chamber; and combustion gases produced by said propellant grain when nozzle means to exhaust said combustion gases for free ignited, said first nozzle means having vblow-out means fiight constructed and arranged to exhaust angularly to which remain therein until said grain is ignited, the the longitudinal axis of said elongated combustion cham other end of said motor tube being closed, a booster tube 30 ber to produce a component of thrust passing through disposed Within said perforation, said booster tube having said seat. an igniter at the end thereof adjacent the closed end of 3. A piston assembly as described in claim 2 wherein said motor tube, and having at the other end thereof said aperture means comprises a plurality of apertures second nozzle means for exhausting combustion gases, and the means for opening the aperture means comprises said second nozzle means being adjacent said iirst nozzle 35 said second propellant being formed from strip propel means, said first and second nozzle means being con lant having a predetermined initial web thickness mounted structed and arranged to exhaust angularly to the lon over the apertures on the inside surface of .the inner gitudinal axis of the` motor tube to produce a com chamber, the strip propellant being adapted to act as a ponent of thrust passing through the seat, said booster membrane in closing the apertures to passage of com tube having a plurality of apertures along the length 40 bustion gases produced in the chamber into the main com thereof and extending through its walls, said apertures bustion chamber at the start of ejection and being further being closed with strips of propellant fixed to the inner adapted to rupture at a predetermined time after the surface of said booster tube, said igniter being operable start of ejection as a result of the decrease in membrane to ignite said strips of propellant for producing combus web thickness and strength resulting from burning. tion gases, said combustion gases produced by said strips 45 4. A piston assembly as described in claim 3 wherein of propellant being operable to cause said rocket motor assembly to travel out of the launcher tube, said strips of propellant being further adapted to form a membrane said nozzle means comprises a nozzle communicating with the main chamber for exhausting combustion gases pro~ duced therein, said nozzle initially closed by a blow-out over said apertures and prevent combustion gases pro plug to prevent premature ignition of the main chamber propellant by entry of gases produced in the inner cham ber therethrough, and a nozzle communicating with the inner chamber for exhausting the gases produced therein and exhausting said portion of the gases produced in the main chamber which exhaust through the inner chamber. duced by said strips from passing through said apertures until said rocket motor assembly has moved to a prede termined position along said launching tube at which po sition said strips of propellant are adapted to rupture and permit hot combustion gases produced by said strips to pass through said apertures, said hot gases passing 55 through said apertures being operable to effect the igni tion of said propellant grain, the combustion gases pro duced by said propellant grain being operable to pro duce thrust to project the seat away from the aircraft after said rocket motor assembly leaves the launcher tube, 60 a portion of said combustion gases produced by said pro pellant grain after ignited adapted to pass through the apertures in the booster tube whereby gas is simultaneous References Cited in the file of this patent UNITED STATES PATENTS 2,434,652 2,724,237 2,814,179 2,869,463 2,877,504 2,900,150 Hickman ____________ .__ Hickman ____________ _Edelman et al. _______ _.. McKnight ___________ _.. Fox ________________ _... Hirt et al. ___________ .__ Jan. 20, 1948 Nov. 22, 1955 Nov. Ian. Mar. Aug. 26, 20, 17, 18, 1957 1959 1959 1959 ly exhausted through said first and second nozzle means. 2. In an aircraft seat ejection system having a launcher 65 OTHER REFERENCES tube secured to the aircraft, a piston assembly slideably Aviation Week Magazine, Nov. 12, 1956, vol. 65, No. engaged in the launcher tube and a seat fixedly attached 20, pages 71, 72, 74 and 77 relied upon.