US3164955A - Turbo compressor drive for jet power plant - Google Patents

Turbo compressor drive for jet power plant Download PDF

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US3164955A
US3164955A US768497A US76849758A US3164955A US 3164955 A US3164955 A US 3164955A US 768497 A US768497 A US 768497A US 76849758 A US76849758 A US 76849758A US 3164955 A US3164955 A US 3164955A
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fluid
turbine
jet
air
heat exchange
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George H Garraway
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    • FMECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F02COMBUSTION ENGINES; HOT-GAS OR COMBUSTION-PRODUCT ENGINE PLANTS
    • F02KJET-PROPULSION PLANTS
    • F02K7/00Plants in which the working fluid is used in a jet only, i.e. the plants not having a turbine or other engine driving a compressor or a ducted fan; Control thereof
    • F02K7/08Plants in which the working fluid is used in a jet only, i.e. the plants not having a turbine or other engine driving a compressor or a ducted fan; Control thereof the jet being continuous
    • FMECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F01MACHINES OR ENGINES IN GENERAL; ENGINE PLANTS IN GENERAL; STEAM ENGINES
    • F01KSTEAM ENGINE PLANTS; STEAM ACCUMULATORS; ENGINE PLANTS NOT OTHERWISE PROVIDED FOR; ENGINES USING SPECIAL WORKING FLUIDS OR CYCLES
    • F01K23/00Plants characterised by more than one engine delivering power external to the plant, the engines being driven by different fluids
    • F01K23/02Plants characterised by more than one engine delivering power external to the plant, the engines being driven by different fluids the engine cycles being thermally coupled
    • F01K23/06Plants characterised by more than one engine delivering power external to the plant, the engines being driven by different fluids the engine cycles being thermally coupled combustion heat from one cycle heating the fluid in another cycle
    • FMECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F01MACHINES OR ENGINES IN GENERAL; ENGINE PLANTS IN GENERAL; STEAM ENGINES
    • F01KSTEAM ENGINE PLANTS; STEAM ACCUMULATORS; ENGINE PLANTS NOT OTHERWISE PROVIDED FOR; ENGINES USING SPECIAL WORKING FLUIDS OR CYCLES
    • F01K23/00Plants characterised by more than one engine delivering power external to the plant, the engines being driven by different fluids
    • F01K23/12Plants characterised by more than one engine delivering power external to the plant, the engines being driven by different fluids the engines being mechanically coupled
    • F01K23/16Plants characterised by more than one engine delivering power external to the plant, the engines being driven by different fluids the engines being mechanically coupled all the engines being turbines
    • FMECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F02COMBUSTION ENGINES; HOT-GAS OR COMBUSTION-PRODUCT ENGINE PLANTS
    • F02KJET-PROPULSION PLANTS
    • F02K3/00Plants including a gas turbine driving a compressor or a ducted fan
    • FMECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F02COMBUSTION ENGINES; HOT-GAS OR COMBUSTION-PRODUCT ENGINE PLANTS
    • F02KJET-PROPULSION PLANTS
    • F02K3/00Plants including a gas turbine driving a compressor or a ducted fan
    • F02K3/08Plants including a gas turbine driving a compressor or a ducted fan with supplementary heating of the working fluid; Control thereof
    • 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
    • Y02T50/60Efficient propulsion technologies, e.g. for aircraft

Description

Jan. 12, 1965 G. H. GARRAWAY 3,164,955
TURBO COMPRESSOR DRIVE FOR JET POWER PLANT Filed Oct. 20, 1958 2 Sheets-Sheet 1 JNVENTOR. @fome h. nmomr BY mff@ United States Patent() 3,164,955 TURB@ CMPRESSOR DRIVE FR .llE'I PQWER PLANT George H. Garraway, 8 Holley Lane, Darien, Coun. Filed st. 2li, 1958, Ser. No. 768,497 11 Claims. (Cl. @tl-35.3)
This invention has to do with a jet propulsion power plant having a mechanical air compressor driven by av turbine, and provides a novel system for driving the turbine. In such power plants, which develop a jet thrust by rearward discharge of an expanding stream of gas (commonly gaseous combustion products) and in which air for combustion or for the jet is brought to a high pressure by a rotary compressor, with or without initial compression in a ram or diffuser at the air inlet, it has been common practice to employ some part or all of the gaseous products of the heating step as the driving duid for the turbine which drives the compressor, the gas discharged from the turbine being delivered to the jet nozzle at a super atmospheric pressure. In proposals for using nuclear energy in such jet propulsion power plants, the step of air-fuel combustion is replaced by an air heating step and the heated air, pre-compressed, is to form the jet. For the present purpose, the two are equivalent and are referred to generically as the step of providing a hot gaseous jet fluid.
This practice of using some or all of the high temperature jet fluid to drive the turbine has imposed a limitation on the heating or combustion process because the temperature of the heated air or of the combustion products passing through the turbine must be limited to a value tolerable by the metal of the turbine. This limitation requires that one or more or all of the factors aifecting the temperature of the gaseous jet fluid be limited, such as the degree of compression of the air or (in a combustion system) the fuel/air ratio or the kind of fuel. There has been a diiculty in finding the best compromise of these factors while Working Within the imposed limitation upon the attained temperature. Such practices sometimes lower the eiciency of combustion yand always decrease the propulsive force of thefjet because of the limitation upon the pressure and temperature of the gaseous fluid delivered to the jet. Further, the use of the gaseous jet fluid as the turbine drive uid has involved the diversion of a large part of the thus limited energy potential of that fluid. The effort to minimize the etect of these limitations has been directed to improvements in turbine and compressor design to increase eiciency, and to improvement of metals to permit higher gas temperatures through the turbine; but a serious ternperature limitation has remained and the equipment improvements have involved undesirable Weight and cost.
Another line of attack on the problem has been the practice called afterburning in power plants using combustion products as the jet fluid, that is, the burning of additional fuel in the products of combustion after their expansion through the turbine, utilizing excess air present in the products from the primary combustion. This gives an increase in propulsive effort, but because the fuel is burned at a point of reduced pressure the gain obtained is small and the cost in additional fuel is great. This practice therefore is used only for short periods suchas during take-off of a jet plane or to give increased acceleration under emergency conditions of iiight. Its high fuel requirement prevents its regular use because the addiproving the operation of such a jet power plant by provid- I ing a Way of driving the turbine of the turbocompressor 3,164,955 Patented Jan. `12, 1965 which permits the attainment of higher temperatures and pressures in the air heating or combustion step of the power plant, and consequently a greater jet thrust, with either a decrease in initial load or an increased speed or distance of travel for the same initial load, or both, all in comparison with a system employing the gaseous jet fluid itself as the turbine drive iluid.
In my system, the turbine of the turbo-compressor unit is notv driven by gaseous jet fluid but rather by a's'epa'rate uid which is in ra gaseous state at least when passing through the turbine and which is circulated in a closed loop including a heat exchange stage at which the circulating uid is heated by indirect heat exchange with a hot jet fluid at one or more points in` the path from air inlet to jet discharge. The hot fluid providing the heat for the separate turbine iiuid may be the hot jet fluid at the heating or combustion `step and at the jet nozzle, or the hot air passing from the ram diffuser to the rotary compressor, ory both.
. In this system, the selections of the fuel, the degree of compression and such factors as the fuel/air ratio for combustion, and also of the temperatures and pressures attained in the gaseous jet fluid, are independent of any limitations imposed by the turbine, and therefore are suchV as to give a greater pressure drop through the jet nozzie and a greater propulsive force from the jet, often Vwith improved heating'or combustion efficiency. Also, it
becomes unnecessary to use such a heavymulti-stage rotary compressor, and a simpler and lighter compressor may be used which effects such a reduction of weight and cost as tomore than justify its tolerably lower eiiiciency of mechanical compression. v Other advantages, and particular preferred features, are Vdescribed below in the description of an illustrative form shown schematically in FIGURE l of the accompanying drawing. FIGURE 2 shows a similar system schematically, including a further feature of novelty.
The illustrative system of the FIG. l diagram is one in which the gaseous jet fluid consists of products Vof fuelwhich in turn delivers to the jet nozzle 14. When nuclear,- energy is used, either alone or with anair-fuel combusf tion system, the combustor 13 is either replaced or sup'- plemented byYV a heater in which the pre-compressed air from lll is heated by the source of nuclear energy.- The 'construction and arrangement of-these several parts is not a part of the present novelty and may take any of various forms. When air-fuel combustionis used, fuel in fluid state is supplied to the combustor i3 by a line lwwhich may` include one or more heat exchangers for pre-heating and preferably vaporizing the fuel before it is delivered to the combustor.
The gas turbine` l5 of the turbo-compressor unit is drivenby a separate duid, separate, that is, from the jet fluid, and this separate turbine uid is circulated in av closed loop indicated by the heavy broken line 20. l
Considered in the direction of ow from turbine discharge 21 back to the turbine inlet 26, this loop includes a heat exchange unit 22, prefehably serving also as a condenser vwhen the circulating liuid is Aa vaporizable liquid, a storage vessel or accumulator 23, a pressure-increasing fluid circulator 24, a heat exchange stage indicated generally as 25 (described in greater detailbelow), and the optional fuel pre-heater for super-heater 36.
The preferred turbine drive fluid is a material that is liquid at atmospheric temperature and pressure, that has a relatively low latent heat of vaporization, and that is compatible with the metals of the various components through which it flows in the closed loop Z0. Examples of suitable fluid are dichlorodiuoromethane (CClgFg), having a latent heat of vaporization on the order of 4() Btu. per pound, and fluorinated cyclic ether (CSFmO), having a latent heat of vaporization of 37.5 Btu. per pound, a high density, relatively low viscosity, and a wide spread between its freezing point (-l7l F.) and its boiling point Q-224 F.).
)In the case where such a vaporizable liquid is employed in the loop 2u, the circulator 24 is a pump suitable for pumping a liquid, the heat exchange stage serves to vaporize the circulating liquid, by indirect heat exchange with gaseous jet fluid flowing to or through the jet nozzle, and the turbine uid is then in a gaseous state when it reaches the turbine inlet 25. Preferably there is no condensation within the turbine and the lluid discharging from the turbine at 21, still in gaseous state, flows to the accumulator 23 by way of heat exchanger 22 in which .it is cooled and preferably condensed to liquid state by indirect heat exchange, such as with inflowing relatively cold fuel supplied through the line 16. T he pump 24 then draws liquid from the accumulator 23, delivers the liquid at increased pressure, and maintains the circulation in the closed loop in which this cycle of vaporization, pressure drop through the turbine and condensation are accomplished.
The heat exchange stage 25, where such a drive fluid is vaporized, involves preferably two steps of indirect heat exchange between circulating drive lluid and the gaseous jet uid. The first step, capable of vaporizing the drive fluid if desired, is at the jet nozzle 14 and cumbustor (or heater) 13 and is made possible by casings 30 and 30a which surrround each of these units so as to leave annular spaces '31 and 31a through which the circulating turbine fluid can be passed as indicated by the spiral lines 32 and 32a. Suitable baffles are provided in these annular spaces 31 and 31a to cause the lluid to circulate while moving axially. The turbine fluid may be passed also to a second heat exchanger 34, serving as a super-heater and in the form of a coil located in the path of the gaseous jet fluid flowing from the combustor or heater 13 to the jet nozzle 14. This heat exchanger 34 is shown in the function-al schematic as lying between the combustor (heater) 13 and the jet nozzle 14. In the structure itself, in which units 13 and 14 may be physically united, this heat exchange coil may be within either unit. In such a system, the turbine fluid is preferably vaporized in the annular spaces 31 and 31a. It is desirable to provide a valve-controlled by-pass 55 through which the circulating turbine fluid in whole or in part may pass also to the heat exchanger 34 after passing through the annular spaces 31 and 31a surrounding the jet nozzle and combustor; and of course various alternative arrangements are possible to vary the order in which the fluid passes through these heat exchange steps, all of this detail being a matter of engineering choice in the particular case.
Inflowing fuel, pre-heated and preferably vaporized in the heat exchanger 22, by indirect heat exchange with the turbine drive fluid discharging from the turbine, passes from the heat exchange 22 to a secondary and optional heat exchanger 35 and thence to the combustor 13. This heat exchanger 36 is on the inlet side of the turbine and serves to eect indirect heat exchange between the preheated fuel and the high temperature circulating turbine drive fluid flowing from the heat exchange stage 25.
Alternatively, the turbine drive uid may be one which is in gaseous state at fall times. In this case, the circulator 24 is a gas compressor and the heat exchanger 22, at the turbine discharge, serves only as a cooler for the circulating uid and a pre-heater for the fuel, without condensing the circulating fluid.
In speaking of the turbine drive fluid as circulating in a closed loop, l distinguish from a once-through system having no re-circulation; but I do not exclude, although I prefer to avoid, some diversion or discard or loss of the circulating fluid and a corresponding make-up. The degree of closure of the physical elements of the loop need be only such as to prevent unintended losses. For example, the accumulator 23 may be vented to the atmosphere in a system in which the fluid is condensed after the turbine. Such a vent, or an equivalent means of establishing a pressure datum, is preferable.
The turbine 15 may be coaxialwith the compressor 1l or may be on a separate axis, with any suitable mechanical driving connection between the two in either case, various designs for this purpose being well known; The turbine i5 may serve also to drive the circulator24, as indicated by the dashdot line 37, and may also drive other auxiliaries of the jet power plant. The storage vessel or accumulator 23, which serves to'maintain a small uid head on the inlet of the circulator 24% and to absorb any fluctuations in the closed loop 2t), is of relatively lightweight construction because the circulating drive fluid, when it reaches this vessel, is at relatively low pressure. In some instances a portion of the conduit may serve as the accumulator.
One of the advantages of this system is that is permits greater freedom as to the spacing and relative location of the combustor (heater), the turbine and the jet nozzle, whereas in systems using gaseous jet fluid to drive the turbine it is almost imperative that the turbine be close to the combustor or heater 13. This greater freedom in physical design is of especial value where a multiple jet nozzle is used.
The ram or diffuser 10 for effecting an initial compres sion of the air taken in at the air inlet is not essential to the basic system but is highly desirable. lts potential advantages are more fullyvrealized in the present system because of the possibility of conductingV the combustion or heating in a way to attain higher'prcssures inthe hot gaseous jet fluid.
In this system, requiring noresort to excess combustion air to hold down the temperature of the combustion products in a combustion system, the ratio of fuel to air can be substantially stoichiometric, and the combustor may be so designed on known principles as to yield combustion products of higher temperature and pressure than in the past, since the combustion products in passing to and through the jet nozzle cornein contact only with static surfaces which can be cooled. A greater pressure drop through the jet nozzle is therefore possible and a consequent greater propulsive effort or thrust.
A further contribution to the gain in thrust is available because in this system, with its higher temperature and pressure of the gaseous jet huid, a much smaller fraction of the energy content of that uid is diverted to turbine operation than is necessary when the jet fluid itself is used as the turbine drive fluid. Operation of the jet nozzle with jet Huid of higher temperature than in the past is facilitated especially when the circulating -drive Huid for the turbine is vaporized in the annular spaces 31 and 31a around the nozzle and combustor or heater 13, since this vaporization provides a good means of cooling the walls 0f these high temperature units.
If the system of the present invention is used in a jet power plant that develops the same jet thrust as a comparable plant using gaseous jet fluid to drive the turbine, the advantage of the system is realized as a large reduction of the rate of air ilow through the compressor, combustor (heater) and jet nozzle. Operation of aircraft at higher altitudes becomes possible with this system, both because the increase in jet thrust permits easier attainment of high altitudes and because, with less air available at the high altitude, the system permits attainment of a suflicient thrust to maintain llight. All of these advantages are expressed as an increase of thrust per unit weight of air flow per second due to the higher attainable pressure and temperature of the gaseous jet fluid.
The increased thrust attainable with this system makes it feasible to employ an air-breathing iet power plant as the first stage of a multi-stage missile or space vehicle. The engine weight and fuel load required for a jet power plant so used, and employing this system, is less than that for a liquid or solid fuel rocket, so that the advantage can be realized as either a faster acceleration or a longer firststage fiight, or both, or simply as a reduction in weight.
The advantage of greater jet thrust per unit of air flow per second appears also in the smaller percentage reduction of thrust with increase in air inlet velocity. As is known, jet thrust is a function of the difference between outlet (jet) velocity and air inlet velocity, so that thrust is reduced as the vehicle attains a higher velocity causing a higher air inlet velocity. The subtractive effect `of inlet velocity, being more nearly a constant, is smaller in terms of the percentage reduction of thrust when the jet velocity is greater, as it can be when the pressure of the 'gaseous jet fluid delivered to the jet nozzle can be increased, as with this system of turbine drive.
FIGURE 2 is a similar schematic flow diagram of a system having all of the elements shown in FIGURE l (except the optional fuel pre-heater 36) in the same relationship as in the system of FIGURE 1, but with certain additional elements now described, the iiow of air, `fuel and turbine drive fluid being as shown by the arrows. He-re again, the illustrative form is one using air-fuel combustion to provi-de the jet fluid, but the same principle is applicable if the combustor is replaced or supplemented by a heater for pre-compressed air with nuclearfuel as the source of heat.
The principal addition in this system of FIGURE 2 is a heat exchanger 4f) effecting indirect heat exchange between (a) the iniiowing air discharging at elevated pres,- sure and temperature from the ram diffuser 10, and (b) the turbine drive fluid circulating in loop 20.
A second added element is a regenerator Si), an optional unit, in which heat exchange (preferably indirect) is effected between the circulating turbine drive fluid as it is discharged from the turbine at 21 and the same fiuid as it is delivered at lower temperature from the pressureincreasing circulator 24 after having been cooled (and condensed, when a vaporizable fluid) in the fuel pre-heating heat exchanger 22. This lowers the temperature of the turbine fluid prior to full `condensation and therefore permits or facilitates condensation, but without loss of heat from the cycle as a whole since the `heat is transferred to a point in the cycle (beyond the point of condensation) where a high temperature is desirable.
A third additional element is shown at 6i), representing a use of the turbine drive fluid to pick up heat from one or more structural parts of the aircraft which become overheated by air friction at very high speeds such as the wing leading edges. This provides additional heat for the turbine drive fluid, but its primary utility is the cooling action on the airframe. This is effected by incorporating a portion of the turbine fluid flow conduit in the portion of the airframe to be cooled.
At very high speeds of a jet propelled air-breathing aircraft, the ram diffuser lil effects a great increase in pressure of the air passing through it, by the well-known ram action. This involves a great increase in air temperature as well, a matter which imposes a limitation upon the speed at which the aircraft can operate well because of the adverse effect upon the rotary compressor 1l. Because of the increase in jet propulsive force attainable with my system of turbine drive, with its removal of the limit upon the tolerable temperature of the compressor (11) output, air speeds are attainable which would involve an excessively high temperature in the output of a ram diffuser. diffuser outlet, and ahead of the rotary mechanical compressor 1li, I provide a heat sink which lowers the By placing the heat exchanger 4u at the ram limiting fractor, and permits the use of additional ram` diffuser capacity as by the provision of multiple diusers.
The heat thus taken from the ram air serves to provide 4G can be varied.
a part of the heat for the turbine drive fluid. Under some conditions, this heat exchange at 46 can vaporize and even superheat the turbine drive iiuid, leaving for heat exchange (25) at the jet nozzle and combuster (heater) 13 only a further super-heating function; but the distribution of heating load between 'thetwo heat exchangersZS and Provision is made also for a bleed of a part of the thus cooled ram air by line 41, controlled by valve 42 or by other means providing expansion and incidental further cooling. This line 41 discharges to the atmosphere either directly or after use of the air as a coolant or refrigerant in one or more parts of the aircraft requiring cooling for comfort or protection, or after use of it as a further heat sink to dissipate-heat from the turbine fluid at any suitable point in its cycle, such as at the turbine discharge. This alternative isl shown in FIG. 2, wherein the bled air passing through line 41 is fed into heat exchanger. 61 and by opening valve 63, turbine fluid Amay pass through line 62 and through exchanger 61,
where it is cooled by heat exchange with the bled ram air. The turbine fluid may then be returned to line 2t)V through line 64 and valve 65. The bled ram air leaving exchanger 61 may then'be discharged to the atmosphere.
The heat exchanger 40 can be by-passed when desired, either by by-passing the ram air around the exchanger 40 or by by-passing the turbine drive fluid aroundit through line 43,'such aresultbeing effected Aby opening valve 44 and closing valve 45. This is for operation at lower aircraft speeds where the heat effect of the ram action is less and the heat requirement of the turbine drive fluid is less. In that event, the turbine drive fluid is less. VIn that event, the turbine drive fluid delivered from circulator 24- is passed directly to the jet nozzle heat exchanger 25 as in FIGURE 1,.,instead of first being heated in exchanger '40. f Y
Considerations of the engineering design of a particular jet vpower plant may make it desirable to pass the turbine drive uid first through heat exchanger Z5 and then through heat exchanger 4t), and l include that variant within. my invention.
This system permit operation of an aircraft at higher speeds and higher altitudes than have been attainable before. For such high speed, high altitude operation, the turbine drive iiuid is preferabl yone having properties such as those of the fluorinated cyclic ether described above, which is an 'available product of Minnesota Min-V ing and Manufacturing Company, .sold under the name of inert Liquid FAC-75. Y Y
Having regard to the two disclosed forms of my system, it will be seen'that basic toboth is the .provision in a jet propulsion power system as described, of a separate turbine drive uid circulated in a Vclosed loop to and ,through'the turbine with pressure increase between the turbine dschargev and turbine inlet, and a heating of this` fluid (with or without'vaporization) Yby Vtransfer Vto it, through indirect heat exchange means, of heat provided by a relatively hot jet fluid, either the ram air discharging y from the diffuser (i.e., ahead of the rotary mechanical compressor and en route to the combustor or heater lf3.) or from the hot gaseous jet fluid proper, or from both in sequence. For this purpose I treat the ram air as a jet uid since it either serves directly as the jet fluid d proper after further compression and heating or as a reagent in the combustion process when the gaseous jet iiuid proper takes the form of products of air-fuel cornbustion. In the case where air-fuel combustion is used, my system also provides a pre-heat, and preferably a vaporization, of the fuel for combustion by transfer to it, also by indirect heat exchange, of heat contained in the turbine drive uid as it is discharged from the turbine, this heat exchange serving to cool the circulating uid and to condense it when it is a vaporizable liquid as preferred. When the jet uid proper takes the form of heated precompressed air (heated, for example, by nuclear fuel) the turbine i'luid is preferably cooled, and condensed when vaporizable, by indirect heat exchange with cooled ram air bled off as by line 41 in FIGURE 2 as well as by the action of the regenerator 50.
I claim:
1. A Ijet propulsion power system comprising an air inlet, a mechanical compressor for inlet air, a gas turbine for driving the compressor, a ram diffuser for receiving and compressing inlet air ahead of said mechanical compressor, a hot gaseous jet fluid generating chamber receiving compressed air from the compressor, a jet nozzle receiving hot gaseous jet iluid from the generating chamber and forming a propulsive jet, means for supplying a Vseparate driving iluid, distinct from said hot gaseous jet huid, to said turbine in a gaseous state which comprises a closed loop ow conduit between the turbine discharge and the turbine inlet, with means therebetween for circulating a contained uid through the turbine and conduit, means lfor effecting indirect heat exchange between said circulating fluid and relatively hot jet iiuid and means for eiecting indirect heat exchange between said circulating turbine drive uid and relatively hot air discharging from said ram diffuser.
2. A system in accordance with claim 1 and further including means for transferring heat to said turbine uid from the gaseous jet fluid at and beyond the generating chamber of hot gaseous jet uid.
3. A system in accordance with claim 1 and further including means for bleeding off ram air after cooling by said heat exchange means for 'use as a heat sink in the said system.
4. In a jet propulsion power system comprising the components of a ram diuser and a turbine driven cornpressor for compressing inlet air, and a het gaseous jet fluid generating chamber receiving air thus compressed and delivering said jet uid to a jet-forming nozzle, wherein the said components provide a ow of jet fluid at elevated temperature from said diffuser to and through said nozzle, the method Iof supplying gaseous driving vtluid lto the turbine which comprises circulating a separate uid distinct from said hot gaseous `jet fluid, in a closed loop to and through said turbine, increasing the pressure of said uid as it passes from turbine discharge to turbine inlet, heating said fluid before return delivery to the turbine by eecting indirect heat exchange between said fluid and said jet cduid of elevated temperature and |between said fluid and ram air discharging from said diffuser.
5. The method 4of claim 4 and which further cornprises bleeding off ram air cooled by said heat exchange and effecting heat exchange between said bled ram air and turbine Ifluid discharging from the turbine to cool the latter.
6. The method of claim 4 and which further comprises effecting an indirect heat exchange between the turbine iiuid and gaseous jet uid at and beyond the generating chamber of the hot gaseous jet fluid.
Cil
7. A jet propulsion power system comprising an air inlet, a mechanical compressor `for inlet air, a gas tur- -bine for driving the compressor, a hot rgaseous jet fluid generating chamber receiving compressed air from the compressor, a jet nozzle receiving hot gaseous jet uid from said generating chamber and forming a propulsive jet, means for supplying a separate driving `fluid, distinct from said hot gaseous `jet duid, to said turbine in a gaseous state which comprises a closed loop flow conduit between the turbine discharge and the turbine inlet, means for circulating a contained iluid through the turbine and conduit and means lfor electing indirect heat exchange between said circulating fluid and relatively hot jet uid within that portion of the system comprising the jet nozzle and thereby cooling said jet nozzle.
8. A jet propulsion power system comprising an air inlet, a mechanical compressor `for inlet air, a `gas turbine for driving the compressor, a hot gaseous jet iuid generating chamber receiving compressed air from the compressor, a jet nozzle receiving hot gaseous jet iluid from said generating chamber and lforming a propulsive jet, means for supplying a separatedriving fluid, distinct from said hot gaseous jet fluid, to said turbine in a gaseous state which comprises a closed loop flow conduit between the turbine discharge and the turbine inlet, means for circulating a contained fluid through the turbine and conduit and means for effecting indirect heat exchange :between said circulating fluid and relatively hot jet iiuid within that portion of the `system comprising the generating chamber and the jet nozzle, and thereby cooling said generating chamber and jet nozzle.
9. A system in accordance with claim 7 in which said heat exchange means effects vaporizat'ion of said circulating fluid and wherein there is disposed a condenser in said closed loop conduit at the discharge side of said turbine to liquefy the circulating fluid after discharge.
1). A system lin accordance with claim 7 in which the hot gaseous jet fluid generating chamber is a combustor with means for delivering fuel -for combustion, and in which there is a second heat exchange means for effecting indirect heat exchange between inilowing fuel and the circulating turbine drive ttluid at the discharge side of the turbine.
11. A system in accordance with claim 7 `in which the hot gaseous jet iiuid generating chamber is a combustor- `with means for delivering fuel for combustion, in which the circulating turbine drive iiuid contained in said loop is a vaporizable liquid adapted to be vaporized "by said heat exchange with :relatively hot jet iluid, and in which there is a condenser at the discharge side of the turbine adapted to liquify the gaseous turbine discharge, with means for passing inowing fuel therethrough as an 1`ndirectly acting condensing medium.
References Cited in the le of this patent UNTED STATES PATENTS 2,154,481 Verkauf Apr. 18, 1939 2,159,758 Diedrich May 23, 1939 2,483,045 Harby Sept. 27, 1949 2,586,025 Godfrey Feb. 19, 1952 2,820,599 Ackeret et al. Jan. 21, 1958 2,955,422 `Peterson Oct. 11, 1960 2,970,437 Anderson Feb. 7, 1961 3,016,694 Howarth et al. Jan. 16, 1962 FOREIGN PATENTS,
679,007 Great Britain Sept. 10, 1952

Claims (1)

  1. 4. IN A JET PROPULSION POWER SYSTEM COMPRISING THE COMPONENTS OF A RAM DIFFUSER AND A TURBINE DRIVEN COMPRESSOR FOR COMPRESSING INLET AIR, THUS COMPRESSED FLUID GENERATING CHAMBER RECEIVING AIR THUS COMPRESSED AND DELIVERING SAID JET FLUID TO A JET-FORMING NOZZLE, WHEREIN THE SAID COMPONENTS PROVIDE A FLOW OF JET FLUID AT ELEVATED TEMPERATURE FROM SAID DIFFUSER TO AND THROUGH SAID NOZZLE, THE METHOD OF SUPPLYING GASEOUS DRIVING FLUID TO THE TURBINE WHICH COMPRISES CIRCULATING A SEPARATE FLUID DISTINCT FROM SAID HOT GASEOUS JET FLUID, IN A CLOSED LOOP TO AND THROUGH SAID TURBINE, INCREASING THE PRESSURE OF SAID FLUID AS IT PASSES FROM TURBINE DISCHARGE TO TURBINE INLET, HEATING SAID FLUID BEFORE RETURN DELIVERY TO THE TURBINE BY EFFECTING INDIRECT HEAT EXCHANGE BETWEEN SAID FLUID AND SAID JET FLUID OF ELEVATED TEMPERATURE AND BEING SAID FLUID AND RAM AIR DISCHARGING FROM SAID DIFFUSER.
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Cited By (20)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
US3516254A (en) * 1967-09-11 1970-06-23 United Aircraft Corp Closed-loop rocket propellant cycle
US3722220A (en) * 1963-11-20 1973-03-27 Texaco Inc Reaction propulsion engine and method of operation
US3739581A (en) * 1972-01-19 1973-06-19 E Talmor Method and apparatus for providing jet propelled vehicles with a heat sink
US4811556A (en) * 1986-10-14 1989-03-14 General Electric Company Multiple-propellant air vehicle and propulsion system
US4817890A (en) * 1986-10-14 1989-04-04 General Electric Company Multiple-propellant air vehicle and propulsion system
US4835959A (en) * 1986-10-14 1989-06-06 General Electric Company Multiple-propellant air vehicle and propulsion system
US4840025A (en) * 1986-10-14 1989-06-20 General Electric Company Multiple-propellant air vehicle and propulsion system
US4841723A (en) * 1986-10-14 1989-06-27 General Electric Company Multiple-propellant air vehicle and propulsion system
FR2640322A1 (en) * 1988-12-09 1990-06-15 Europ Propulsion Rocket motor, or combined motor for a space vehicle with an essentially closed auxiliary hydraulic circuit
US5247792A (en) * 1992-07-27 1993-09-28 General Electric Company Reducing thermal deposits in propulsion systems
US5805973A (en) * 1991-03-25 1998-09-08 General Electric Company Coated articles and method for the prevention of fuel thermal degradation deposits
US5891584A (en) * 1991-03-25 1999-04-06 General Electric Company Coated article for hot hydrocarbon fluid and method of preventing fuel thermal degradation deposits
US20110179766A1 (en) * 2009-10-27 2011-07-28 Fly Steam, LLC Heat recovery system
US20110232298A1 (en) * 2010-03-23 2011-09-29 General Electric Company System and method for cooling gas turbine components
WO2011131493A2 (en) * 2010-04-20 2011-10-27 Rolls-Royce Plc Air breathing reaction propulsion engine
RU2555609C2 (en) * 2013-08-15 2015-07-10 Федеральное государственное бюджетное образовательное учреждение высшего профессионального образования Самарский государственный технический университет Combined cycle cooling unit operating method and device for its implementation
US20190047722A1 (en) * 2017-08-09 2019-02-14 Hs Marston Aerospace Limited Fuel tank inerting system
RU2722519C1 (en) * 2019-07-09 2020-06-01 Федеральное государственное унитарное предприятие "Научно-производственный центр автоматики и приборостроения имени академика Н.А. Пилюгина" (ФГУП "НПЦАП") Stabilization method of structurally unstable carrier rocket oscillators
US10941706B2 (en) 2018-02-13 2021-03-09 General Electric Company Closed cycle heat engine for a gas turbine engine
US11015534B2 (en) 2018-11-28 2021-05-25 General Electric Company Thermal management system

Citations (9)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
US2154481A (en) * 1933-01-09 1939-04-18 Herpen & Vorkauf Power plant
US2159758A (en) * 1936-09-24 1939-05-23 Gen Electric Power plant
US2483045A (en) * 1945-09-24 1949-09-27 Harold D Harby Jet engine, including a combustion chamber to which gaseous fuel is delivered under pressure
US2586425A (en) * 1947-10-13 1952-02-19 American Dairies Inc Method of iodinating proteins
GB679007A (en) * 1950-03-13 1952-09-10 Bbc Brown Boveri & Cie Thermal power plant
US2820559A (en) * 1956-11-30 1958-01-21 John H Armitage Auxiliary cargo vehicle for trucking vehicles
US2955422A (en) * 1958-09-19 1960-10-11 Adolphe C Peterson Power turbine and pressure fluid means
US2970437A (en) * 1956-02-28 1961-02-07 Thompson Ramo Wooldridge Inc High temperature pumping system with variable speed pump and refrigeration by-product
US3016694A (en) * 1958-05-13 1962-01-16 Westinghouse Electric Corp Combination turbojet and ramjet engine

Patent Citations (9)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
US2154481A (en) * 1933-01-09 1939-04-18 Herpen & Vorkauf Power plant
US2159758A (en) * 1936-09-24 1939-05-23 Gen Electric Power plant
US2483045A (en) * 1945-09-24 1949-09-27 Harold D Harby Jet engine, including a combustion chamber to which gaseous fuel is delivered under pressure
US2586425A (en) * 1947-10-13 1952-02-19 American Dairies Inc Method of iodinating proteins
GB679007A (en) * 1950-03-13 1952-09-10 Bbc Brown Boveri & Cie Thermal power plant
US2970437A (en) * 1956-02-28 1961-02-07 Thompson Ramo Wooldridge Inc High temperature pumping system with variable speed pump and refrigeration by-product
US2820559A (en) * 1956-11-30 1958-01-21 John H Armitage Auxiliary cargo vehicle for trucking vehicles
US3016694A (en) * 1958-05-13 1962-01-16 Westinghouse Electric Corp Combination turbojet and ramjet engine
US2955422A (en) * 1958-09-19 1960-10-11 Adolphe C Peterson Power turbine and pressure fluid means

Cited By (22)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
US3722220A (en) * 1963-11-20 1973-03-27 Texaco Inc Reaction propulsion engine and method of operation
US3516254A (en) * 1967-09-11 1970-06-23 United Aircraft Corp Closed-loop rocket propellant cycle
US3739581A (en) * 1972-01-19 1973-06-19 E Talmor Method and apparatus for providing jet propelled vehicles with a heat sink
US4840025A (en) * 1986-10-14 1989-06-20 General Electric Company Multiple-propellant air vehicle and propulsion system
US4817890A (en) * 1986-10-14 1989-04-04 General Electric Company Multiple-propellant air vehicle and propulsion system
US4835959A (en) * 1986-10-14 1989-06-06 General Electric Company Multiple-propellant air vehicle and propulsion system
US4841723A (en) * 1986-10-14 1989-06-27 General Electric Company Multiple-propellant air vehicle and propulsion system
US4811556A (en) * 1986-10-14 1989-03-14 General Electric Company Multiple-propellant air vehicle and propulsion system
FR2640322A1 (en) * 1988-12-09 1990-06-15 Europ Propulsion Rocket motor, or combined motor for a space vehicle with an essentially closed auxiliary hydraulic circuit
US5805973A (en) * 1991-03-25 1998-09-08 General Electric Company Coated articles and method for the prevention of fuel thermal degradation deposits
US5891584A (en) * 1991-03-25 1999-04-06 General Electric Company Coated article for hot hydrocarbon fluid and method of preventing fuel thermal degradation deposits
US5247792A (en) * 1992-07-27 1993-09-28 General Electric Company Reducing thermal deposits in propulsion systems
US20110179766A1 (en) * 2009-10-27 2011-07-28 Fly Steam, LLC Heat recovery system
US20110232298A1 (en) * 2010-03-23 2011-09-29 General Electric Company System and method for cooling gas turbine components
WO2011131493A2 (en) * 2010-04-20 2011-10-27 Rolls-Royce Plc Air breathing reaction propulsion engine
WO2011131493A3 (en) * 2010-04-20 2012-03-08 Rolls-Royce Plc Air breathing reaction propulsion engine
RU2555609C2 (en) * 2013-08-15 2015-07-10 Федеральное государственное бюджетное образовательное учреждение высшего профессионального образования Самарский государственный технический университет Combined cycle cooling unit operating method and device for its implementation
US20190047722A1 (en) * 2017-08-09 2019-02-14 Hs Marston Aerospace Limited Fuel tank inerting system
US10518895B2 (en) * 2017-08-09 2019-12-31 Hs Marston Aerospace Limited Fuel tank inerting system
US10941706B2 (en) 2018-02-13 2021-03-09 General Electric Company Closed cycle heat engine for a gas turbine engine
US11015534B2 (en) 2018-11-28 2021-05-25 General Electric Company Thermal management system
RU2722519C1 (en) * 2019-07-09 2020-06-01 Федеральное государственное унитарное предприятие "Научно-производственный центр автоматики и приборостроения имени академика Н.А. Пилюгина" (ФГУП "НПЦАП") Stabilization method of structurally unstable carrier rocket oscillators

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