(12) United States Patent (10) Patent No.: US 6,619,031 B1

Transcription

1 USOO B1 (12) United State Patent (10) Patent No.: Balepin (45) Date of Patent: Sep. 16, 2003 (54) MULTI-MODE MULTI-PROPELLANT OTHER PUBLICATIONS LOUID ROCKET ENGINE O NGIN D. Huzel and D Huang, Modern Engineering for Deign of (76) Inventor: Vladimir V. Balepin, 3350 Keokuk St., Liquid-Propellant Rocket Engine' Vol 147, AIAA Serie Butte, MT (US) Progre in Atronautic and Aeronautic pp. 35, 36 (1992). G. Oate, Aerothermodynamic of Ga Turbine and Rocket (*) Notice: Subject to any diclaimer, the term of thi Propulion. AIAA Education Serie, p. 52 (1988). patent i extended or adjuted under 35 H. Hirakoo et al. A Concept of Lace for Space Plane to U.S.C. 154(b) by 0 day. Earth Orbit, Int. J. Hydrogen Energy, Vol 15, No. 7, pp (1990). 21) Appl. No.: 09/842,535 Y. Miki, et wu wu-ko al., Advanced Scram-Lace Svtem y Concept p for p pp (21) Appl. No.: 09/842, Single-Stage-to-Orbit Space Plane. IAF , pp 1-6 (22) Filed: Apr. 25, 2001 International Atronautical Federation (1991). V. Balepin, et al., Lightweight Low Cot Klin Cycle Related U.S. Application Data Derivative for a Small Reuable Launcher, (60) Proviional application No. 60/200,129, filed on Apr. 27, AIAA , pp 1-5, AIAA (Nov. 1-4, 1999) M. Togawa, et al., A Concept of Lace for SSTO Space (51) Int. Cl.... F02K5/00 Plane AIAA , pp 1-6, AIAA (1991). (52) U.S. Cl /246 * cited by examiner (58) Field of Search... 60/246, 257, 260 Primary Examiner Ehud Gartenberg (56) Reference Cited (74) Attorney, Agent, or Firm-Ralph F. Crandell U.S. PATENT DOCUMENTS (57) ABSTRACT 3,747,339 A 7/1973 Wolf et al. A multi-mode multi-propellant rocket engine capable of 3,756,024 A * 9/1973 Gay... 60/204 operating in a plurality of Selected mode. 3,768,254 A * 10/1973 Stuart... 60/204 3,775,977. A 12/1973 Builder et al. GF(Kx + 1)C 3, A * 3/1974 Haumann et al /268 All- = 1 4,073,138 A 2/1978 Beichel O 4,393,039 A 7/1983. Sherman 4,771,600 A 9/1988 Limerick et al. Propellant component may include liquid hydrogen, liquid 4,782,655 A * 11/1988 Weber /772 hydrocarbon, liquid oxygen, liquid fluorine, and liquid air A 5/1989 Martin A 1f1990 Hulin 60/204 The liquid oxygen and the liquid air are Stored in Separate 5025,623 A 6/1991 Rideral tank are mixed in a dedicated mixer prior to their injection A 9/1991 Bond.r into the combution chamber. 5,101,622 A 4/1992 Bond 5,154,051 A * 10/1992 Mouritzen... 60/257 3 Claim, 5 Drawing Sheet ar zzzzzzzzzzzzzzzzzzzzzzzzzzzzzzzzzzzzzzzzzzzzzzzzzzzzzzzzzzzzzzzzzzzzzzzzzzzzzzzzzzz!

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4 U.S. Patent Sep. 16, 2003 Sheet 3 of 5 (%) NIV9 NOILOVH-I GWOT1)\\/d CNln TWO MODES Fig. 3 THREE MODES (%) EWnTOA HE HONOVTI TWO MODES Fig. 4 THREE PROPUSVE MODES

5 U.S. Patent Sep. 16, 2003 Sheet 4 of 5 M - 10% OXDZER PUMP CONTROL ARLIOEFACTION RATIO Fig. 5 L CO - O 2 CD l 4000 CO oo O ooo ALL-ROCKET 3000 O FLIGHT MACH NUMBER Fig. 6

6 U.S. Patent Sep. 16, 2003 Sheet 5 of 5 13 H (f) O C H l 2 H CC - 12 O FLIGHT MACH NUMBER Fig ALL-ROCKET 90 CD I - 70 a. - DRYWEIGHT TAKEOFF WEIGHT 50 O INITIAL AIR-TO-HYDROGENRATIO Kao Fig. 8

7 1 MULT-MODE MULTI-PROPELLANT LIQUID ROCKET ENGINE CROSS-REFERENCE TO RELATED APPLICATION Thi application claim the benefit of Proviional Appli cation Ser. No. 60/200,129, filed Apr. 27, 2000, by Vladimir V. Balepin, for LIQUID AIR AUGMENTED ROCKET ENGINE, the dicloure of which i incorporated herein by thi reference. BACKGROUND OF THE INVENTION 1. Field of the Invention The preent invention relate to liquid propellant rocket engine and particularly to multi-mode multi-propellant Single Stage earth to orbit or Suborbital rocket engine. 2. Decription of the Prior Art Rocket propulion for earth to orbit launch vehicle i currently the only practical choice. Known rocket engine, however, operate at efficiencie that are far from the optimum, particularly for Single Stage earth to orbit opera tion. While many olution for the enhancement of liquid fuel rocket engine have been propoed, few of them have been implemented. Mot innovation lead to ignificant deign complication and cot increae which offet their potential benefit. A a reult, principle dicovered early in the 20th century, Such a for example multitage rocket, Still provide the main bai for modern rocket launcher. It i well known from rocket theory that launcher efficiency can be increaed if high thrut, dene propellant, moderate Specific impule engine are ued at low altitude for the initial acceleration, and high Specific impule, lower thrut engine are ued for high altitude acceleration and orbiting. Liquid propellant combination embodying oxidizer and fuel which have been conidered are liquid oxygen(lox)/ keroene and liquid oxygen (LOX)/liquid hydrogen(lh2) engine; LOX/methane and LOX/LH2 engine; and olid propellant motor and LOX/LH2 engine. Rocket engine uing Such fuel combination have been built and teted, and mot of them are workhore for modern launch buinee. The ue of different pair of fuel for different hardware unit eentially require a multitage configuration of rocket engine. Single-tage-to-orbit (SSTO) launcher are not feaible becaue idling group of engine mut be carried. Multitage configuration further reult in high launch cot and prevent introduction of reuability. The ue a fuel efficient LOX/LH2 engine a the only thruter for SSTO launcher i unlikely to be ucceful becaue uch an engine i not efficient when ued a a low altitude thruter. U.S. Pat. No. 4,771,599 and 4,771,600 dicloe tripro pellant rocket engine utilizing a tripropellant fuel Sytem in which the propellant are oxygen, hydrogen, and a hydro carbon. Such an engine will produce the thrut neceary for initial acceleration of Significant payload into low earth orbit. Thi engine i referred to a a booter or high thrut low altitude engine providing for initial acceleration of the launcher, and it ha only one operational mode. It i not advantageou to ue thi engine for high-peed acceleration and orbiting becaue it ha a lower pecific impule than the conventional LOX/LH2 engine. U.S. Pat. No. 4,831,818 dicloe a dual-fuel, dual-mode Single Stage rocket engine for earth to orbit operation. The fuel are a high Specific impule fuel Such a liquid hydrogen, and a high denity-impule fuel Such a liquid methane. Flow of the fuel i aid to be controlled by the fuel pump. The fuel are ued to cool the nozzle. The fuel are mixed uptream of the nozzle cooling jacket, and the fuel mixture i fed to the cooling jacket. A method i decribed wherein the mixture of fuel i varied to provide a progre Sively less dene mixture while providing thrut. U.S. Pat. No. 5,101,622 dicloe a rocket engine capable of operating in two propulive mode for near earth and low earth orbit operation. It decribe a firt mode in which the external atmophere i the Source of oxidizer for the fuel. At a high Mach number the engine change to a Second mode which i that of a conventional high performance rocket engine uing liquid oxygen carried on the vehicle to oxidize a liquid hydrogen fuel. The engine i Said to ue common hardware including a liquid hydrogen pump and a combu tor nozzle aembly. The mechanim required to match the working fuel and oxidizer flow in both propulive mode i not dicloed. The engine further include Several turbocom pressor to compre air to a delivery pressure of Several hundred bar, a Serie of heat exchanger, and turbopump, all of which make the engine complicated and expenive to produce and operate and it ma prohibitively high. A liquid air cycle engine, or LACE, i another example of the propulion were one of the propellant component, in thi cae oxidizer, can be changed during flight. Liquid oxygen ued on the main acceleration mode can be com pletely or partially replaced by liquefied ambient air during the firt propulive mode beginning at the initial launch or take off through acceleration from Sea level atmopheric condition to moderate Speed and altitude. Such an engine i hown and decribed in H. Hirakoo, A Concept of LACE for Space Plane to Earth Orbit, Int. J. Hydrogen Energy, Vol. 15, No. 7, pp , at p. 499, There i a clear intention in the LACE concept to maximize the air condenation ratio in an effort to achieve maximum pecific impule or Ip. No real attention i paid, however, to complication to the engine reulting from the neceary additional pump, plumbing, valve, etc. AS a reult, the LACE how inadequate performance gain and/or prohibitively complicated and heavy deign. None of the known LACE decription Sugget the mechanim to match ga flow through the nozzle throat in both the combined and the rocket mode. Fuel Storage Sytem for rocket engine are hown in U.S. Pat. No. 5,804,760, and U.S. Pat. No. 5,705,771. The preent invention can be baed upon exiting rocket engine uing an expander cycle (RL10 of Pratt& Whitney), ga generator cycle (J2 of Boeing-Rocketdyne), tap-off cycle (J2S, RS2000 of Boeing-Rocketdyne), or a taged combution topping cycle (SSME of Boeing-Rocketdyne) known in the art. Variou example of rocket engine can be found in D. Huzel and D. Huang, Modern Engineering for Deign of Liquid-Propellant Rocket Engine, Volume 147 of AIAA Serie ProgreSS in Atronautic and Aeronautic, page 35, 36, (1992). DESCRIPTION OF THE DRAWINGS FIG. 1 i a diagrammatic cross-sectional view of a multi propellant multi-mode rocket engine (MPLRE) embodiment of the preent invention. FIG. 2 i a diagrammatic cross-sectional view of a multi propellant multi-mode liquid air augmented rocket engine (LAARE) embodiment of the preent invention. FIG. 3 i a bar graph howing the projected percent payload fraction gain for different propulive mode com pared to a baic expander type rocket engine. FIG. 4 i a bar graph howing the projected percent launcher Volume for different propulive mode compared to a baic expander type rocket engine.

8 3 FIG. 5 i a graph howing the theoretical relationhip of the oxidizer/hydrogen ratio to the air liquefaction ratio for Selected oxygen/hydrogen ratio. FIG. 6 i a graph howing the theoretical relationhip of the Specific impule to the flight Mach number a a function of variou air liquefaction ratios. FIG. 7 i a graph howing the theoretical relationhip of relative thrut to the flight Mach number a a function of variou air liquefaction ratios. FIG. 8 i a graph howing a theoretical comparion of the relative percentage of rocket weight to the initial air to hydrogen ratio for an all-rocket launcher and for a liquid air augmented rocket engine. DESCRIPTION OF THE INVENTION The preent invention i embodied in a new, novel and unobviou multi-propellant, multi-mode-rocket engine 20 a hown in FIG. 1, and in the method of operation thereof. The rocket engine 20 i formed by a combutor 21 having a hell 22 defining a cylindrical combutor chamber 24 opening at one end through a throat 25 into a wide nozzle 26. At it other end the chamber 22 Support an injector head 28 through which propellant fuel and oxidizer component are introduced into the combutor chamber 24 in which they are ignited and burn to produce exhaut gae to provide the deired thrut. The combutor 21 i provided with an external cooling jacket or hell 29 defining a cooling paage 30 adapted to receive one of the fuel component Such a liquid hydrogen a a coolant for the nozzle 26 and combutor 21. Propellant component in the form of fuel and oxidizer are fed to the injector head from Storage tank therefor. Referring to FIG. 1, propellant component Storage tank include a liquid hydrogen tank 32, a liquid oxygen tank 33, a liquid hydrocarbon tank 34, and a Supplemental oxidizer tank 35 for a Supplemental oxidizer Such a liquid fluorine. For cooling the combutor 21, hydrogen fuel i fed from the tank 32 through a conduit 36 to a turbine driven hydrogen pump 38 and thence through a conduit 39 having a main hydrogen control valve 40 to the cooling chamber paage 30 define by the cooling jacket 29. From the cooling jacket 29 the liquid hydrogen, now warmed by the combutor, flow through a conduit 41 and by expanding drive a pump turbine 42. From the turbine 42 the hydrogen flow through a conduit 44 to the engine injector head 28. To control the turbine peed, a bypa valve 45 allow warmed liquid hydrogen to bypa the turbine 42, thereby controlling the amount of hydrogen flowing through the turbine 42. For driving a econdary pump 46 to feed hydrocarbon fuel from the hydrocarbon Supply tank 34 through a control valve 47 to the engine, a econdary turbine 48 i alo driven by expanding a portion of the hydrogen a determined by a control valve 49 in the hydrogen upply conduit 50. Liquid oxygen i fed from the oxygen Supply tank 33 through a control valve 52, conduit 54, an oxidizer pump 55 driven by the main turbine 42, then through a main oxidizer control valve 56 and conduit 58 into the injector head 28. When the rocket vehicle reache a high altitude and leave the Earth' atmophere, launch efficiency can be increaed by uing propellant component that create toxicity rik at lower altitude. Such component can be metallic additive to the fuel or more efficient oxidizer, for example liquid fluorine. Phae of the flight above atmopheric can ue a fluorine-oxygen mixture in the ratio of about 50 percent each. Thi mixture i dener than the liquid oxygen and can be tored under the ame cryogenic condition. To thi end, the fluorine Storage tank 35 i connected to the main oxidizer upply line 58 through a liquid fluorine control valve 59. The oxygen and fluorine control valve 52, 59 are adjuted during flight to provide the deired ratio of oxidizer. The fluorine and oxygen are preferably Stored a the mixture in one tank thermally integrated with the tank of the liquid OXygen. The injector head 28 feed metered amount of fuel and oxidizer which are burned in the combutor chamber to produce hot gae that then flow through the engine throat and nozzle and are ejected to produce the deired thrut. The thrut produced by the burning gae propel the rocket engine and vehicle, and by the ue of the variou control valve for controlling of the flow of fuel and oxidant a decribed below, the optimum thrut performance can be maintained. Another embodiment of the preent invention i hown in FIG. 2. In decribing thi embodiment, imilar reference numeral will be ued with the ditinguihing uffix a. Thi modification comprie a rocket engine 20a of the foregoing character and an aociated air cooling and liq uefaction unit 60 for producing liquid air at atmopheric altitude to augment the oxidizer, conventionally liquid oxygen. The rocket engine 20a i contructed with a com butor 21a defining a combutor chamber 24a opening through a throat 25a into a wide exhaut nozzle 26a. A propellant fuel, Such a liquid hydrogen, i Supplied to the combutor injector head 28a from a fuel tank 32a through a control valve 4.0a and combutor jacket 29a and cooling chamber 30a by a fuel pump 38a driven by a expanded hydrogen turbine 42a. A Supplemental fuel Such a a hydro carbon i fed to the injector head 28a from a Supply tank thereof 34a through a control valve 47a by a pump 46a driven by a turbine 48a. A propellant oxidizer, Such a liquid oxygen, i fed to the combutor injector head 28a from a upply tank 33a through a control valve 52a by an oxidizer pump 55a. At higher altitude, liquid fluorine oxidizer i fed from a upply tank thereof 35a through a control valve 59a by the pump 55a. The pump 38a and 55a are driven by a main turbine 42a operatively connected thereto and powered by expanding a portion of the liquid hydrogen fuel. The air cooling and liquefaction unit 60 i formed by a liquefaction chamber 61 having an air inlet 62. Liquid oxygen i introduced into the incoming air for initial cooling and moiture freezing-out purpoe through an oxidizer injection Sytem or manifold 64. Liquid oxygen flow to the manifold from the main oxygen line 58a through a conduit 65 and control valve 66 in the main line 58a. The mixed air and oxidizer i further cooled and liquefied in the chamber 61 by contact with a heat exchanger and condener 68 in the chamber 61. A mixer 69 in the main conduit 58a receive liquid and Saturated air from the liquefaction unit 60 through a conduit 70 and mixe it with the liquid oxidizer uch a liquid oxygen from the oxidizer Supply tank 33. Becaue the mixer 69 ha a low Suction head, a low preure turbopump 71 may be utilized to feed the mixed oxidizer from the mixer 69 to the oxidizer pump 55a in order to prevent cavitation therein. A control valve 72 may be provided in the liquid air conduit 70 to provide for liquefied air mode operation and Subequent tranition of the engine to an all-rocket mode. The engine 20 i able to operate on both a liquefied air cycle mode and all-rocket mode. The former mode i characterized by ue of a liquid air and oxygen mixture a an oxidizer. In the all-rocket mode only liquid oxygen i ued a the oxidizer. When operating in the liquid air cycle, air captured through air inlet 62 i cooled and partially liquefied

9 S in the condener 68 cooled by the liquid hydrogen fuel a a coolant. In order to prevent heat exchanger performance deterioration a a reult of icing, the incoming air i cooled prior to entering the condener 68 to a temperature below the water triple point (273.15K) by injection of liquid oxygen from the oxidizer injection manifold 64. Air from the heat exchanger and condener 68 i in Saturated condition with a liquid ma content of more than 80%. Liquefaction i accomplihed in the mixer 69 where cold Saturated air meet the higher flow rate of on-board OXygen which can be ubcooled (55K) to liquefy more air. The liquid air and liquid oxygen product of the mixer 69 i a lightly ubcooled liquid oxidizer that provide cavitation-free operation of the oxidizer pump 55a. The low-preure turbopump 71 i alo ued for obtaining better anti-cavitation characteritic in the oxidizer feeding Sytem. Liquid hydrogen from the heat exchanger and condener 68 flow to the combutor 26 through the combutor cooling jacket 29a. Liquid oxidizer from the oxidizer pump 55a i upplied to the combutor 21a of the rocket engine 20a along with liquid hydrogen fuel pumped by pump 38a driven by driving turbine 42a. Combution product are expanded through the nozzle 26a to generate thrut. After the initial acceleration of the engine, when the humidity of the atmopheric air cannot caue precooler icing, the valve 66 top liquid oxidizer flow into the oxidizer injection Sytem 64. A reaonable peed of tranition to an all-rocket mode correpond to Mach=6-7. Operation of the airbreath ing Sytem above thi peed i not beneficial becaue of cooling requirement. At the tranition Mach number, about mach , air liquefaction i ceaed and the engine 20 operate a a pure rocket engine with liquid hydrogen fuel and liquid oxygen oxidizer. The operation of the engine whether the modification hown in FIG. 1 or the modification hown in FIG. 2 a decribed above i controlled by a computer 80. The com puter receive combutor ga temperature and preure data from a temperature Senor 81 and a pressure Senor 82 in the combutor chamber 24. The flow rate of propellant compo nent i meaured by flowmeter 84, 85, 86, 87, 88 and 89 meauring the flow of liquid oxygen, liquid fluorine, liquid hydrocarbon fuel, liquid hydrogen fuel, liquefied air, and liquefied oxidizer fed to the air liquifaction unit 60, repec tively. All control valve are connected to and their operation i controlled by the computer 80. The computer calculate the optimum ma flow rate of the propellant component, the characteritic exhaut Velocity of the exhaut gae produced in the combutor according to the ga temperature and preure, the compoition of the propellant component and other neceary parameter, and control the flow of propellant component to the injector head 28 a decribed to achieve the deired performance characteritic. FIGS. 3 and 4 how the benefit of multipropellant rocket engine application to SSTO launcher compared to a con ventional LOX/LH2 liquid rocket engine. FIG.3 how thi benefit in the form of the payload fraction gain to the low earth orbit of 407 km, which fraction i given a a percentage of the launcher gro take-off weight. It i een from the FIG. 3 that when two propulive mode are employed (LOX/ LH2/keroene combination from SLS to 28 km), payload fraction gain count for 1.4%. When a third mode i added (FLOX/LH2 at the altitude above 150 km), payload fraction gain increae to 1.79%. It hould be noted that the exiting launcher provide payload fraction to the low earth orbit in the vicinity of 2-3% of the gro take-off weight. FIG. 4 how that both two- and three-propulive mode engine provide Significant Volume reduction of the launcher due to hydrogen fraction reduction. The Volume of the two con idered launcher are repectively 76.1 and 73.5% of the baic launcher with ingle mode LOX/LH2 propulion. Thi ignificant volume reduction provide up to 20% of the vehicle drag reduction that, in turn, increae launcher efficiency. Engine thrut-to-weight reduction for SLS con dition in the conidered example count for 40% compared to the baic LOX/LH2 engine. In the cae of a liquid air augmented rocket engine or LAARE, the following flight cenario from take-off to earth orbit or Suborbital condition occur: 1) Initial launch or take off and acceleration from Sea level condition to moderate hypersonic Speed, uually about Mach The rocket engine, which conventionally operate with a liquid oxygen and liquid hydrogen or LOX/LH2 propellant ytem, utilize liquefied ambient air added to LOX to increae thrut and Specific impule during take-off and initial acceleration to mod erate hypersonic Speed. Hydrocarbon fuel can alo be ued in thi mode for additional efficiency. 2) Major part of acceleration and acent. The rocket engine operate in it primary deign mode with a LOX/LH2 propellant producing moderate thrut and high Specific-impule during the major part of the acceleration, from about Mach toward orbital or Suborbital Speed. Final acceleration can be completed with or without high energy fuel additive to the LOX/ LH2 propellant. FIG. 5 how the oxidizer/hydrogen ratio (K+KA) in the LAARE mode a a function of the air liquefaction ratio K. where Ka=0 correpond to the all-rocket mode, for different oxygen/hydrogen ratio (K= ) in the rocket mode. ASSuming that the oxygen/hydrogen mixture ratio in the Second or rocket mode i K=6.0 and the initial air lique faction ratio i Ka=2.0, a vertical arrow line i drawn from the air liquefaction ratio axi to the line correponding to K=6.0. Next, a horizontal arrow line i drawn to the oxidizer/hydrogen ratio axi. The obtained value of the oxidizer/hydrogen ratio (K+KA)=6.42 in the LAARE mode effectively matche the Selected rocket mode K=6.0. ASSuming that the hydrogen flow rate i the Same in both operational mode in thi example, the oxidizer flow rate are proportional to the indicated ratio number, i.e., 6 in the Second or rocket mode and 6.42 in the firt or LAARE mode. Thi mean that the flow rate through the oxidizer tur bopump in thi example i jut 7% higher in the LAARE operational mode. Thi number i definitely within the control range of modern turbopump. In the engine hown in FIG. 2 incorporating the LAARE Cycle, it i imple to increae turbopump power in the LAARE mode, becaue hydrogen, which i the turbopump driver, i additionally heated in thi mode in the heat exchanger/condener a compared to the all-rocket mode. The dotted line in FIG. 5 correpond to +10% and +15% oxidizer pump flow rate control. The LAARE Cycle engine accordingly doe not require additional pump, turbine or compreor or exotic and complicated mean to increae the air liquefaction ratio, a in the prior art Liquid Air Cycle Engine (LACE) Cycle. See, for example, U.S. Pat. No. 4,393,039, and U.S. Pat. No. 5,025,623. High-preure pump and their driver are ued in both the LAARE and all-rocket mode. Thi reult in a Subtantial increae in the engine thrut-to-weight ratio compared to prior art engine. LAARE Cycle benefit are poible becaue of the unique combination of it parameter. FIGS. 6 and 7 preent a projected comparion of the etimated Specific impule and relative thrut for the LAARE Cycle engine with different air

10 7 liquefaction ratio and pure rocket LH2/LOX propulion. It i etimated that LAARE Cycle engine' Specific impule will be higher, for example, at ea level, an etimated 30% higher, than that of a conventional rocket engine. Subtantial thrut deterioration along the trajectory i a typical weakne of the air-breathing accelerator, which reult in engine overize. In the cae of a LAARE or air liquefaction Cycle engine, etimated thrut i high under Sea level condition and i projected to be higher and nearly contant during acceleration, a hown in FIG. 7. The combination of thee two parameter provide a favorable etimated effective Specific impule, which along with exceptional engine thrut-to-weight ratio, produce a high launcher efficiency. The LAARE Cycle engine make feaible ytem that are not feaible with all-rocket propulion for Small- or mid-ize reuable SSTO launcher. A LAARE Cycle launcher may create a new market for on-demand, Small payload launch Service. Additionally, it may boot Space commerce activitie including Space manufacturing. The LAARE Cycle i alo an attractive propulion option for both a Suborbital global-reach vehicle and a pace Station reupply vehicle. A LAARE Cycle engine application to a Small vertical takeoff SSTO launcher wa conidered in Ref. 3 (V. Balepin and P. Hendrick, Lightweight Low Cot KLIN Cycle Rerivative for a Small Reuable Launcher, ALAA Technical Paper , 1999). FIG. 8 preent a projected comparion of a LAARE Cycle launcher to an all-rocket launcher in term of relative gro takeoff weight (GTOW) and dry weight, with corre ponding parameter of all-rocket launcher taken a 100%, a a function of the initial air liquefaction ratio K.A. According to FIG. 8, the GTOW and dry weight of a BANTAM-cla launcher, launching a 330 pound (1b) pay load to 220 nautical mile (nmi) orbit, utilizing the LAARE Cycle, could be reduced by 45% and 30%, repectively, a compared to the all-rocket launcher. The bet launcher efficiency correpond to Ka= Several improvement can be incorporated into the LAARE Cycle engine baed on the expander cycle rocket engine. The LAARE Cycle working process in the combined cycle mode i accompanied by "enthalpy injection' from incoming air into the hydrogen fuel. The hydrogen tempera ture uptream of the combutor cooling jacket i Kelvin (K) higher, depending on the air liquefaction ratio in thi mode a compared to the rocket mode, which afford a Significant increae in the turbine power and combutor pressure and flow rate. The rocket engine i deigned with a combutor having the requiite fixed or variable nozzle area, and the appro priate propellant fuel and oxidizer for producing the deired thrut. The fuel combination, flow rate and ratio are Selected to provide a high denity-impule fuel during the firt mode and a high Specific impule during the Second mode of operation. The Speed of the rocket in flight and variou controllable engine parameter are Sened by Senor and meter. The nozzle area, the flow and ratios of the propellant component, and other variable parameter are controlled by appropriate mechanim and control valve and are maintained during flight by a Suitable control circuit which may include a computer for receiving the data from the rocket and rocket engine and controlling the rocket' flight. While illutrative embodiment of the preent invention have been hown in the drawing and decribed above in coniderable detail, it hould be undertood that there i no intention to limit the invention to the Specific form di cloed. On the contrary the intention i to cover all modification, alternative contruction, equivalent, meth od and ue falling within the Spirit and Scope of the invention a expreed in the appended claim. What i claimed i: 1. A multi-mode multi-propellant rocket engine capable of operating in a plurality of Selected mode compriing, in combination, a combutor chamber opening through a throat into a wide nozzle, a propellant injector head in Said combutor chamber; a propellant fuel Supply tank containing liquid hydrogen; a propellant oxidizer Supply tank containing liquid oxy gen, a jacket covering Said combutor chamber and nozzle and defining there with a heat exchange chamber; a hydrogen pump for Supplying liquid hydrogen fuel from Said hydrogen Supply tank through Said jacket heat exchanger to Said injector head; a control valve for controlling the Supply of liquid hydro gen to Said injector head; an air liquefaction chamber for receiving and liquefying ar, a mixer for receiving liquefied air from Said air liquefac tion chamber and liquid oxygen from Said oxygen Supply tank and mixing the ame; an oxidizer pump for Supplying liquid oxygen from Said Oxidizer Supply tank to Said injector head, for Supplying a mixture of liquid oxygen and liquid air to Said injector head and for Supplying a mixture of liquid oxygen and liquid air to Said air liquefaction chamber; a Spray manifold in Said air liquefaction chamber for receiving aid liquid air and liquid oxygen mixture and Spraying the ame into the air being liquefied; a heat exchanger in Said air liquefaction chamber for receiving liquid hydrogen from Said liquid hydrogen Supply pump, Said liquid hydrogen heat exchanger cooling Said air to effect liquefaction thereof; a control valve for controlling the Supply of liquefied air and injected oxygen to Said mixer, an oxygen pump for Supplying liquid oxygen and liquid air oxidizer from Said mixer to Said injector head; a control valve for controlling the Supply of liquid oxygen and liquid air to Said injector head; and a turbine for operatively receiving a Supply of expanded hydrogen for driving Said hydrogen pump and Said oxygen pump; the operation of Said control Valve determining and controlling the ratio of propel lant fuel and oxidizer component in each mode of operation of the rocket engine. 2. A multi-mode multi-propellant rocket engine capable of operating in a plurality of Selected mode compriing, in combination, a combutor chamber opening through a throat into a wide nozzle, a propellant injector head in Said combutor chamber; a propellant fuel Supply tank containing liquid hydrogen; a propellant fuel Supply tank containing liquid hydrocar bon; a propellant oxidizer Supply tank containing liquid oxy gen, a jacket covering Said combutor chamber and nozzle and defining there with a heat exchange chamber; a hydrogen pump for Supplying liquid hydrogen fuel from Said hydrogen Supply tank through Said jacket heat exchanger to Said injector head;

11 a control valve for controlling the Supply of liquid hydro gen to Said injector head; a hydrocarbon pump for Supplying liquid hydrocarbon fuel from Said hydrocarbon tank to Said injector head; a hydrogen driven turbine controllably driving Said hydro carbon pump; an air liquefaction chamber for receiving and liquefying air; a mixer for receiving liquefied air from Said air liquefac tion chamber and liquid oxygen from Said oxygen Supply tank and mixing the ame; an oxidizer pump for Supplying liquid oxygen from Said oxidizer Supply tank to Said injector head, for Supplying a mixture of liquid oxygen and liquid air to Said injector head and for Supplying a mixture of liquid oxygen and liquid air to Said air liquefaction chamber; a Spray manifold in Said air liquefaction chamber for receiving Said liquid air and liquid oxygen mixture and Spraying the ame into the air being liquefied; a heat exchanger in Said air liquefaction chamber for receiving liquid hydrogen from Said liquid hydrogen Supply pump, Said liquid hydrogen heat exchanger cooling Said air to effect liquefaction thereof; a control valve for controlling the Supply of liquefied air and injected oxygen to Said mixer, an oxygen pump for Supplying liquid oxygen and liquid air oxidizer from Said mixer to Said injector head; a control valve for controlling the Supply of liquid oxygen and liquid air to Said injector head; and a turbine for operatively receiving a Supply of expanded hydrogen for driving Said hydrogen pump and Said oxygen pump; the operation of Said control valve determining and controlling the ratio of propellant fuel and oxidizer component in each mode of operation of the rocket engine. 3. A rocket engine a defined in claim 1 or 2 further compriing Senor in Said combutor for monitoring the temperature and preure of the gae therein; flowmeter for monitoring the flow of each propellant component; and a computer operative in repone to the temperature and pressure of Said gae, the flow rate of Said propellant component, and the chemical compoition of Said propel lant component for controlling Said control valve to con trol the Supply of propellant component to Said injector head in each mode of operation of the rocket engine. k k k k k