Abstract:
A multi-mode multi-propellant rocket engine capable of operating in a plurality of selected modes.                G   f          (       K   X     +   1     )            C   *           A   *          P   0         =   1                           
     Propellant components may include liquid hydrogen, liquid hydrocarbon, liquid oxygen, liquid fluorine, and liquid air. The liquid oxygen and the liquid air are stored in separate tanks are mixed in a dedicated mixer prior to their injection into the combustion chamber.

Description:
CROSS-REFERENCE TO RELATED APPLICATION 
     This application claims the benefit of Provisional Application Ser. No. 60/200,129, filed Apr. 27, 2000, by Vladimir V. Balepin, for LIQUID AIR AUGMENTED ROCKET ENGINE, the disclosure of which is incorporated herein by this reference. 
    
    
     BACKGROUND OF THE INVENTION 
     1. Field of the Invention 
     The present invention relates to liquid propellant rocket engines and particularly to multi-mode multi-propellant single stage earth to orbit or suborbital rocket engines. 
     2. Description of the Prior Art 
     Rocket propulsion for earth to orbit launch vehicles is currently the only practical choice. Known rocket engines, however, operate at efficiencies that are far from the optimum, particularly for single stage earth to orbit operation. While many solutions for the enhancement of liquid fuel rocket engines have been proposed, few of them have been implemented. Most innovations lead to significant design complications and cost increases which offset their potential benefits. As a result, principles discovered early in the 20th century, such as for example multistage rockets, still provide the main basis for modern rocket launchers. It is well known from rocket theory that launcher efficiency can be increased if high thrust, dense propellant, moderate specific impulse engines are used at low altitudes for the initial acceleration, and high specific impulse, lower thrust engines are used for high altitude acceleration and orbiting. Liquid propellant combinations embodying oxidizers and fuels which have been considered are liquid oxygen(LOX)/kerosene and liquid oxygen (LOX)/liquid hydrogen(LH2) engines; LOX/methane and LOX/LH2 engines; and solid propellant motors and LOX/LH2 engines. Rocket engines using such fuel combinations have been built and tested, and most of them are workhorses for modern launch businesses. The use of different pairs of fuels for different hardware units essentially requires a multistage configuration of rocket engines. Single-stage-to-orbit (SSTO) launchers are not feasible because idling groups of engines must be carried. Multistage configurations further result in high launch costs and prevent introduction of reusability. The use a fuel efficient LOX/LH2 engine as the only thruster for SSTO launchers is unlikely to be successful because such an engine is not efficient when used as a low altitude thruster. 
     U.S. Pat. Nos. 4,771,599 and 4,771,600 disclose tripropellant rocket engines utilizing a tripropellant fuel system in which the propellants are oxygen, hydrogen, and a hydrocarbon. Such an engine will produce the thrust necessary for initial acceleration of significant payloads into low earth orbit. This engine is referred to as a booster or high thrust low altitude engine providing for initial acceleration of the launcher, and it has only one operational mode. It is not advantageous to use this engine for high-speed acceleration and orbiting because it has a lower specific impulse than the conventional LOX/LH2 engine. 
     U.S. Pat. No. 4,831,818 discloses a dual-fuel, dual-mode single stage rocket engine for earth to orbit operation. The fuels are a high specific impulse fuel such as liquid hydrogen, and a high density-impulse fuel such as liquid methane. Flow of the fuels is said to be controlled by the fuel pumps. The fuels are used to cool the nozzle. The fuels are mixed upstream of the nozzle cooling jacket, and the fuel mixture is fed to the cooling jacket. A method is described wherein the mixture of fuels is varied to provide a progressively less dense mixture while providing thrust. 
     U.S. Pat. No. 5,101,622 discloses a rocket engine capable of operating in two propulsive modes for near earth and low earth orbit operations. It describes a first mode in which the external atmosphere is the source of oxidizer for the fuel. At a high Mach number the engine changes to a second mode which is that of a conventional high performance rocket engine using liquid oxygen carried on the vehicle to oxidize a liquid hydrogen fuel. The engine is said to use common hardware including a liquid hydrogen pump and a combustor nozzle assembly. The mechanism required to match the working fuel and oxidizer flow in both propulsive modes is not disclosed. The engine further includes several turbocompressors to compress air to a delivery pressure of several hundred bars, a series of heat exchangers, and turbopumps, all of which make the engine complicated and expensive to produce and operate and its mass prohibitively high. 
     A liquid air cycle engine, or LACE, is another example of the propulsion were one of the propellant components, in this case oxidizer, can be changed during flight. Liquid oxygen used on the main acceleration mode can be completely or partially replaced by liquefied ambient air during the first propulsive mode beginning at the initial launch or take off through acceleration from sea level atmospheric conditions to moderate speed and altitude. Such an engine is shown and described in H. Hirakoso, “A Concept of LACE for Space Plane to Earth Orbit,” Int. J. Hydrogen Energy, Vol. 15, No. 7, pp. 495-505, at p. 499, 1990. 
     There is a clear intention in the LACE concepts to maximize the air condensation ratio in an effort to achieve maximum specific impulse or Isp. No real attention is paid, however, to complications to the engine resulting from the necessary additional pumps, plumbing, valves, etc. As a result, the LACE shows inadequate performance gain and/or prohibitively complicated and heavy design. None of the known LACE descriptions suggest the mechanism to match gas flow through the nozzle throat in both the combined and the rocket modes. 
     Fuel storage systems for rocket engines are shown in U.S. Pat. No. 5,804,760, and U.S. Pat. No. 5,705,771. 
     The present invention can be based upon existing rocket engines using an expander cycle (RL10 of Pratt&amp; Whitney), gas generator cycle (J2 of Boeing-Rocketdyne), tap-off cycle (J2S, RS2000 of Boeing-Rocketdyne), or a staged combustion topping cycle (SSME of Boeing-Rocketdyne) known in the art. Various examples of rocket engines can be found in D. Huzel and D. Huang, “Modern Engineering for Design of Liquid-Propellant Rocket Engines,” Volume 147 of AIAA Series “Progress in Astronautics and Aeronautics,” pages 35, 36, (1992). 
    
    
     DESCRIPTION OF THE DRAWINGS 
     FIG. 1 is a diagrammatic cross-sectional view of a multi-propellant multi-mode rocket engine (MPLRE) embodiment of the present invention. 
     FIG. 2 is a diagrammatic cross-sectional view of a multi-propellant multi-mode liquid air augmented rocket engine (LAARE) embodiment of the present invention. 
     FIG. 3 is a bar graph showing the projected percent payload fraction gain for different propulsive modes compared to a basic expander type rocket engine. 
     FIG. 4 is a bar graph showing the projected percent launcher volume for different propulsive modes compared to a basic expander type rocket engine. 
     FIG. 5 is a graph showing the theoretical relationship of the oxidizer/hydrogen ratio to the air liquefaction ratio for selected oxygen/hydrogen ratios. 
     FIG. 6 is a graph showing the theoretical relationship of the specific impulse to the flight Mach number as a function of various air liquefaction ratios. 
     FIG. 7 is a graph showing the theoretical relationship of relative thrust to the flight Mach number as a function of various air liquefaction ratios. 
     FIG. 8 is a graph showing a theoretical comparison of the relative percentage of rocket weights 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 present invention is embodied in a new, novel and unobvious multi-propellant, multi-mode-rocket engine  20  as shown in FIG. 1, and in the method of operation thereof. The rocket engine  20  is formed by a combustor  21  having a shell  22  defining a cylindrical combustor chamber  24  opening at one end through a throat  25  into a wide nozzle  26 . At its other end the chamber  22  supports an injector head  28  through which propellant fuel and oxidizer components are introduced into the combustor chamber  24  in which they are ignited and burn to produce exhaust gases to provide the desired thrust. 
     The combustor  21  is provided with an external cooling jacket or shell  29  defining a cooling passage  30  adapted to receive one of the fuel components such as liquid hydrogen as a coolant for the nozzle  26  and combustor  21 . 
     Propellant components in the form of fuels and oxidizers are fed to the injector head from storage tanks therefor. Referring to FIG. 1, propellant component storage tanks 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 as liquid fluorine. 
     For cooling the combustor  21 , hydrogen fuel is 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 passage  30  define by the cooling jacket  29 . From the cooling jacket  29  the liquid hydrogen, now warmed by the combustor, flows through a conduit  41  and by expanding drives a pump turbine  42 . From the turbine  42  the hydrogen flows through a conduit  44  to the engine injector head  28 . To control the turbine speed, a bypass valve  45  allows warmed liquid hydrogen to bypass the turbine  42 , thereby controlling the amount of hydrogen flowing through the turbine  42 . 
     For driving a secondary pump  46  to feed hydrocarbon fuel from the hydrocarbon supply tank  34  through a control valve  47  to the engine, a secondary turbine  48  is also driven by expanding a portion of the hydrogen as determined by a control valve  49  in the hydrogen supply conduit  50 . 
     Liquid oxygen is 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 reaches a high altitude and leaves the Earth&#39;s atmosphere, launch efficiency can be increased by using propellant components that create toxicity risks at lower altitudes. Such components can be metallic additives to the fuel or more efficient oxidizers, for example liquid fluorine. Phases of the flight above atmospheric can use a fluorine-oxygen mixture in the ratio of about 50 percent each. This mixture is denser than the liquid oxygen and can be stored under the same cryogenic conditions. To this end, the fluorine storage tank  35  is connected to the main oxidizer supply line  58  through a liquid fluorine control valve  59 . The oxygen and fluorine control valves  52 ,  59  are adjusted during flight to provide the desired ratio of oxidizers. The fluorine and oxygen are preferably stored as the mixture in one tank thermally integrated with the tank of the liquid oxygen. 
     The injector head  28  feeds metered amounts of fuel and oxidizers which are burned in the combustor chamber to produce hot gases that then flow through the engine throat and nozzle and are ejected to produce the desired thrust. The thrust produced by the burning gases propels the rocket engine and vehicle, and by the use of the various control valves for controlling of the flow of fuels and oxidants as described below, the optimum thrust performance can be maintained. 
     Another embodiment of the present invention is shown in FIG.  2 . In describing this embodiment, similar reference numerals will be used with the distinguishing suffix “a.” This modification comprises a rocket engine  20   a  of the foregoing character and an associated air cooling and liquefaction unit  60  for producing liquid air at atmospheric altitudes to augment the oxidizer, conventionally liquid oxygen. The rocket engine  20   a  is constructed with a combustor  21   a  defining a combustor chamber  24   a  opening through a throat  25   a  into a wide exhaust nozzle  26   a . A propellant fuel, such as liquid hydrogen, is supplied to the combustor injector head  28   a  from a fuel tank  32   a  through a control valve  40   a  and combustor jacket  29   a  and cooling chamber  30   a  by a fuel pump  38   a  driven by a expanded hydrogen turbine  42   a . A supplemental fuel such as a hydrocarbon is fed to the injector head  28   a  from a supply tank thereof  34   a  through a control valve  47   a  by a pump  46   a  driven by a turbine  48   a . A propellant oxidizer, such as liquid oxygen, is fed to the combustor injector head  28   a  from a supply tank  33   a  through a control valve  52   a  by an oxidizer pump  55   a . At higher altitudes, liquid fluorine oxidizer is fed from a supply tank thereof  35   a  through a control valve  59   a  by the pump  55   a . The pumps  38   a  and  55   a  are driven by a main turbine  42   a  operatively connected thereto and powered by expanding a portion of the liquid hydrogen fuel. 
     The air cooling and liquefaction unit  60  is formed by a liquefaction chamber  61  having an air inlet  62 . Liquid oxygen is introduced into the incoming air for initial cooling and moisture freezing-out purposes through an oxidizer injection system or manifold  64 . Liquid oxygen flows to the manifold from the main oxygen line  58   a  through a conduit  65  and control valve  66  in the main line  58   a . The mixed air and oxidizer is further cooled and liquefied in the chamber  61  by contact with a heat exchanger and condenser  68  in the chamber  61 . A mixer  69  in the main conduit  58   a  receives liquid and saturated air from the liquefaction unit  60  through a conduit  70  and mixes it with the liquid oxidizer such as liquid oxygen from the oxidizer supply tank  33 . Because the mixer  69  has a low suction head, a low pressure turbopump  71  may be utilized to feed the mixed oxidizer from the mixer  69  to the oxidizer pump  55   a  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 subsequent transition of the engine to an all-rocket mode. 
     The engine  20  is able to operate on both a liquefied air cycle mode and all-rocket mode. The former mode is characterized by use of a liquid air and oxygen mixture as an oxidizer. In the all-rocket mode only liquid oxygen is used as the oxidizer. When operating in the liquid air cycle, air captured through air inlet  62  is cooled and partially liquefied in the condenser  68  cooled by the liquid hydrogen fuel as a coolant. In order to prevent heat exchanger performance deterioration as a result of icing, the incoming air is cooled prior to entering the condenser  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 condenser  68  is in saturated condition with a liquid mass content of more than 80%. Liquefaction is accomplished in the mixer  69  where cold saturated air meets the higher flow rate of on-board oxygen which can be subcooled (55K) to liquefy more air. The liquid air and liquid oxygen product of the mixer  69  is a slightly subcooled liquid oxidizer that provides cavitation-free operation of the oxidizer pump  55   a . The low-pressure turbopump  71  is also used for obtaining better anti-cavitation characteristics in the oxidizer feeding system. 
     Liquid hydrogen from the heat exchanger and condenser  68  flows to the combustor  26  through the combustor cooling jacket  29   a . Liquid oxidizer from the oxidizer pump  55   a  is supplied to the combustor  21   a  of the rocket engine  20   a  along with liquid hydrogen fuel pumped by pump  38   a  driven by driving turbine  42   a . Combustion products are expanded through the nozzle  26   a  to generate thrust. After the initial acceleration of the engine, when the humidity of the atmospheric air cannot cause precooler icing, the valve  66  stops liquid oxidizer flow into the oxidizer injection system  64 . A reasonable speed of transition to an all-rocket mode corresponds to Mach=6-7. Operation of the airbreathing system above this speed is not beneficial because of cooling requirements. At the transition Mach number, about mach 6.0-6.5, air liquefaction is ceased and the engine  20  operates as a pure rocket engine with liquid hydrogen fuel and liquid oxygen oxidizer. 
     The operation of the engine whether the modification shown in FIG. 1 or the modification shown in FIG. 2 as described above is controlled by a computer  80 . The computer receives combustor gas temperature and pressure data from a temperature sensor  81  and a pressure sensor  82  in the combustor chamber  24 . The flow rate of propellant components is measured by flowmeters  84 ,  85 ,  86 ,  87 ,  88  and  89  measuring 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 , respectively. All control valves are connected to and their operation is controlled by the computer  80 . The computer calculates the optimum mass flow rate of the propellant components, the characteristic exhaust velocity of the exhaust gases produced in the combustor according to the gas temperature and pressure, the composition of the propellant components and other necessary parameters, and controls the flow of propellant components to the injector head  28  as described to achieve the desired performance characteristics. 
     FIGS. 3 and 4 show the benefits of multipropellant rocket engine application to SSTO launcher compared to a conventional LOX/LH2 liquid rocket engine. FIG. 3 shows this benefit in the form of the payload fraction gain to the low earth orbit of 407 km, which fraction is given as a percentage of the launcher gross take-off weight. It is seen from the FIG. 3 that when two propulsive modes are employed (LOX/LH2/kerosene combination from SLS to 28 km), payload fraction gain counts for 1.4%. When a third mode is added (FLOX/LH2 at the altitude above 150 km), payload fraction gain increases to 1.79%. It should be noted that the existing launchers provide payload fraction to the low earth orbit in the vicinity of 2-3% of the gross take-off weight. FIG. 4 shows that both two- and three-propulsive mode engines provide significant volume reduction of the launcher due to hydrogen fraction reduction. The volume of the two considered launchers are respectively 76.1 and 73.5% of the basic launcher with single mode LOX/LH2 propulsion. 
     This significant volume reduction provides up to 20% of the vehicle drag reduction that, in turn, increases launcher efficiency. Engine thrust-to-weight reduction for SLS conditions in the considered example counts for 40% compared to the basic LOX/LH2 engine. 
     In the case of a liquid air augmented rocket engine or LAARE, the following flight scenario from take-off to earth orbit or suborbital conditions occurs: 
     1) Initial launch or take off and acceleration from sea level conditions to moderate hypersonic speed, usually about Mach 6-6.5. The rocket engine, which conventionally operates with a liquid oxygen and liquid hydrogen or LOX/LH2 propellant system, utilizes liquefied ambient air added to LOX to increase thrust and specific impulse during take-off and initial acceleration to moderate hypersonic speed. Hydrocarbon fuel can also be used in this mode for additional efficiency. 
     2) Major part of acceleration and ascent. The rocket engine operates in its primary design mode with a LOX/LH2 propellant producing moderate thrust and high specific-impulse during the major part of the acceleration, from about Mach 6-6.5 towards orbital or suborbital speed. Final acceleration can be completed with or without high energy fuel additive to the LOX/LH2 propellant. 
     FIG. 5 shows the oxidizer/hydrogen ratio (K X +K A ) in the LAARE mode as a function of the air liquefaction ratio K A  where K A =0 corresponds to the all-rocket mode, for different oxygen/hydrogen ratios (K OX =5.0-6.5) in the rocket mode. Assuming that the oxygen/hydrogen mixture ratio in the second or rocket mode is K OX =6.0 and the initial air liquefaction ratio is K A =2.0, a vertical arrow line is drawn from the air liquefaction ratio axis to the line corresponding to K OX =6.0. Next, a horizontal arrow line is drawn to the oxidizer/hydrogen ratio axis. The obtained value of the oxidizer/hydrogen ratio (K X +K A )=6.42 in the LAARE mode effectively matches the selected rocket mode K OX =6.0. 
     Assuming that the hydrogen flow rate is the same in both operational modes in this example, the oxidizer flow rates are proportional to the indicated ratio numbers, i.e., 6 in the second or rocket mode and 6.42 in the first or LAARE mode. This means that the flow rate through the oxidizer turbopump in this example is just 7% higher in the LAARE operational mode. This number is definitely within the control range of modern turbopumps. In the engine shown in FIG. 2 incorporating the LAARE Cycle, it is simple to increase turbopump power in the LAARE mode, because hydrogen, which is the turbopump driver, is additionally heated in this mode in the heat exchanger/condenser as compared to the all-rocket mode. The dotted lines in FIG. 5 correspond to +10% and +15% oxidizer pump flow rate control. The LAARE Cycle engine accordingly does not require additional pumps, turbines or compressors or exotic and complicated means to increase the air liquefaction ratio, as in the prior art Liquid Air Cycle Engine (LACE) Cycles. See, for example, U.S. Pat. No. 4,393,039, and U.S. Pat. No. 5,025,623. High-pressure pumps and their drivers are used in both the LAARE and all-rocket modes. This results in a substantial increase in the engine thrust-to-weight ratio compared to prior art engines. 
     LAARE Cycle benefits are possible because of the unique combination of its parameters. FIGS. 6 and 7 present a projected comparison of the estimated specific impulse and relative thrust for the LAARE Cycle engine with different air liquefaction ratios and pure rocket LH2/LOX propulsion. It is estimated that LAARE Cycle engine&#39;s specific impulse will be higher, for example, at sea level, an estimated 30% higher, than that of a conventional rocket engine. Substantial thrust deterioration along the trajectory is a typical weakness of the air-breathing accelerators, which results in engine oversize. In the case of a LAARE or air liquefaction Cycle engine, estimated thrust is high under sea level conditions and is projected to be higher and nearly constant during acceleration, as shown in FIG.  7 . The combination of these two parameters provides a favorable estimated effective specific impulse, which along with exceptional engine thrust-to-weight ratio, produces a high launcher efficiency. 
     The LAARE Cycle engine makes feasible systems that are not feasible with all-rocket propulsion for small- or mid-size reusable SSTO launchers. A LAARE Cycle launcher may create a new market for on-demand, small payload launch services. Additionally, it may boost space commerce activities including space manufacturing. The LAARE Cycle is also an attractive propulsion option for both a suborbital global-reach vehicle and a space station resupply vehicle. A LAARE Cycle engine application to a small vertical takeoff SSTO launcher was considered in Ref. 3 (V. Balepin and P. Hendrick, “Lightweight Low Cost KLIN Cycle Rerivative for a Small Reusable Launcher,” ALAA Technical Paper 99-4893, 1999). 
     FIG. 8 presents a projected comparison of a LAARE Cycle launcher to an all-rocket launcher in terms of relative gross takeoff weight (GTOW) and dry weight, with corresponding parameters of all-rocket launcher taken as 100%, as a function of the initial air liquefaction ratio K AO . According to FIG. 8, the GTOW and dry weight of a BANTAM-class launcher, launching a 330 pound (lb) payload to 220 nautical miles (nmi) orbit, utilizing the LAARE Cycle, could be reduced by 45% and 30%, respectively, as compared to the all-rocket launcher. The best launcher efficiency corresponds to K AO =2.0-3.0. 
     Several improvements can be incorporated into the LAARE Cycle engine based on the expander cycle rocket engine. The LAARE Cycle working process in the combined cycle mode is accompanied by “enthalpy injection” from incoming air into the hydrogen fuel. The hydrogen temperature upstream of the combustor cooling jacket is 50-80 Kelvin (K) higher, depending on the air liquefaction ratio in this mode as compared to the rocket mode, which affords a significant increase in the turbine power and combustor pressure and flow rate. 
     The rocket engine is designed with a combustor having the requisite fixed or variable nozzle area, and the appropriate propellant fuels and oxidizers for producing the desired thrust. The fuel combinations, flow rates and ratios are selected to provide a high density-impulse fuel during the first mode and a high specific impulse during the second mode of operation. The speed of the rocket in flight and various controllable engine parameters are sensed by sensors and meters. The nozzle area, the flow and ratios of the propellant components, and other variable parameters are controlled by appropriate mechanisms and control valves 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&#39;s flight. 
     While illustrative embodiments of the present invention have been shown in the drawings and described above in considerable detail, it should be understood that there is no intention to limit the invention to the specific forms disclosed. On the contrary the intention is to cover all modifications, alternative constructions, equivalents, methods and uses falling within the spirit and scope of the invention as expressed in the appended claims.