US2694899A - Liquid fuel vaporizing apparatus - Google Patents

Liquid fuel vaporizing apparatus Download PDF

Info

Publication number
US2694899A
US2694899A US167115A US16711550A US2694899A US 2694899 A US2694899 A US 2694899A US 167115 A US167115 A US 167115A US 16711550 A US16711550 A US 16711550A US 2694899 A US2694899 A US 2694899A
Authority
US
United States
Prior art keywords
fuel
nozzles
combustion
starting
heat
Prior art date
Legal status (The legal status is an assumption and is not a legal conclusion. Google has not performed a legal analysis and makes no representation as to the accuracy of the status listed.)
Expired - Lifetime
Application number
US167115A
Inventor
Floyd T Hague
Current Assignee (The listed assignees may be inaccurate. Google has not performed a legal analysis and makes no representation or warranty as to the accuracy of the list.)
CBS Corp
Original Assignee
Westinghouse Electric Corp
Priority date (The priority date is an assumption and is not a legal conclusion. Google has not performed a legal analysis and makes no representation as to the accuracy of the date listed.)
Filing date
Publication date
Application filed by Westinghouse Electric Corp filed Critical Westinghouse Electric Corp
Priority to US167115A priority Critical patent/US2694899A/en
Application granted granted Critical
Publication of US2694899A publication Critical patent/US2694899A/en
Anticipated expiration legal-status Critical
Expired - Lifetime legal-status Critical Current

Links

Images

Classifications

    • FMECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F02COMBUSTION ENGINES; HOT-GAS OR COMBUSTION-PRODUCT ENGINE PLANTS
    • F02CGAS-TURBINE PLANTS; AIR INTAKES FOR JET-PROPULSION PLANTS; CONTROLLING FUEL SUPPLY IN AIR-BREATHING JET-PROPULSION PLANTS
    • F02C7/00Features, components parts, details or accessories, not provided for in, or of interest apart form groups F02C1/00 - F02C6/00; Air intakes for jet-propulsion plants
    • F02C7/22Fuel supply systems
    • F02C7/224Heating fuel before feeding to the burner
    • 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

NOV- 23, 1954 F. T. HAGUE LIQUID FUEL vAPoRIzING APPARATUS Filed June 9, 1950 2 Sheets-Sheet 1 INVENTOR FLOYD T. HGUE f'wvv ATTORNEY lllllllllllllll NOV. 23, 1954 F T HAGUE 2,694,899
LIQUID FUEL VAPORIZING APPARATUS Filed June 9, 1950 2 Sheets-Sheet 2 PRESSURE 20o 3 409 50o soo 'loo soo wlTNEssEs; TEMPERATURE F lNvENToR :H FLoYDKT. HAGUE vAPoRlzATloN CHART 771C FOR MIL-F 5624 FUEL a ATTORN EY vconverted Sinto"y compared to `a gas combustor.
Away under high altitude conditions.
'-'rioydf Tfnague, nrexei iiiii, Pa.; assigner' tu Westinghouse` Electric Corporation; poration of Pennsylvania Apparaten rune' e, 'i'asogffseriai No; 4'167,115
2 Ciaims. (ci. iso-39.14)
VThis inyefntionfrelatesto combustion apparatus 'ofi-the continuous-combustion-itype,-'for fe'xarnple, "such as are used in turbojets,1'propjets'and-1thodyds for propulsion of aircraft, and has' Afor-an object" tolprovideimproved t and *novel-'apparatus' of this character.
vTo obtainfthe rho'st benetfromfany combustiori'proclaws and specifically avoid being handicapped-by'any offthem. The-burning of liquidi and vgaseous 'types of fuels depends "on 'several factors,` not '--the' least "'ofwhich Y is f the characteristiclproperty ofa--liq'uids'- inability to support flame. lLiquids cannot burn; they'must'r'st be b gases or solidsj either of Which 4may urn.
f This statement'v of fact is' generally accepted, but l such acceptancel does not 'carry any1assurancethatiallthe significance of this fact--is fully appreciated. 'The' essentially'necess'ary timefofconv'ersion of liquid 'fuel to 1 gas, lwhich 'must proceed ignition, involves fthe processes of (l) subdivision into lsmall particles (usually called -z itomiza'tion) and (2) theabsorptionof evaporation by-these particles in `order to' evolve 'the of lthe latentv l heat gaseous state which is pre-requisiteto burning. Thus the combustion of liquid fuel inf an atomizingcombustor involves' two-additional time consuming' operations as G'as" cc'nnbustion is lfundamentally `faster than liquid fuel combustion and Jafull understanding of this fact will explain "why' the additional "time delayin ignition-all of the fuel in an atomizing combustor is responsible 'fora `loss in combustion elliciency under certain limiting conditions `which prevail in combustors at Vhigh altitudes.
ALoss of' combustion eticiency means that"the'fin ally evolved gas-is'incompletelycombusted 'as' it leaves -the v combustor. s action is lengthened as the'altitude increases. Inanavia- The time tocomplete the `combustion"'re tion combustor, becauseof itsl geometry, thelzur'ningl gasair `mixture m'ayat `altitude be quenchedl by l'dilution air before the combustion' reaction 1s completed, orinother instances the time delay infevolving' the gaseousfsta'te from the liquid fuel mayresultin 'soine`of the.'V gas-air 'mixture being formed'intheso-called vqueneh section of fthe combustor, where the temperature level'is too lo'wj to fully complete lthe combustion'reactions a'nd 'as a result partially combuste'd ".'gas #leaves the combustor.V 'The downstream distribution patternof atoinized fuel lis a particularly adverse situationiniany high altitude combustor because -the largest "size fuel droplets travel f the farthest before being converted --to -the` gasu state.
`idl'factors which limitthe time that the fn'ally'ev'olved gasis in the complete combustion zonei 'offa' combustor combine toreduce the-combustion etliciency-'inainaior In *ari otherwise perfectly'propo'rtioned liquid fuel 'atomizing"coinbus`t.or,
4these factors 'include (1') lineness of atomizat'io'n, 1particularly at-partialload fuel flows `an'd'atcruising speed, (-2) combustor overall length, '(3) velocity 'of gases throught-he' combustor, (4) distributionpattern of atomized fuel particles, (5) degree f-and type ofnturbulence in the complete combustion'zone and niany less-l important factors.
An evaluation lstudy ofvatomizing 'combustors and'of vaporizing combustorsv indicates-that 'f the atmizii'i'g' aviation combustor -is -handicap'ped `Ifrom attaining "optimum performance by-natural-laws whereas the-samelaws can assist 'a vapori-zingcombustor,v lThe -tir'ried'elay #factors 'of 'downstream distribution of atomiz'ed fuel Aand l"ab- United States Pareti'td 'thousandths of asecond.
^ cruising fuel 'ilow conditions.
'coarser as 'theair pressure level'lwers.
2,694,899 fre-telef! Neff-.1??? 3354 ICC sorption of latent heat of vaporizationbefore'ignition l have'not'beenfound'to beof'suffcient practicalimportance to deserve much consideration eXCepV'in" short "aviation cornb'ust'orsV for"high"alt`itude*operation and 'residencel 'tim'ei-so large'r infact 'thatf under the 'jaltitude conditions 'outlined the efcieiicyofic'ornbust ri 'can be very yadversely aifected. Y 'Low' combustion etiiciency prej'ce'des combustor jblowoutand" all 'existing'combus'tors 'po te conditioncharacterized by "the"ab'sence"of` either tionss'inceithere is no time delay 'in the v combustion chamber.
' and liquid fuel througha' heated vaporizing tube.`
another, liquid fuelalone Was` passed th'ro'ugha heated exchanger in `tlieupper portion 'of the tube. `poor 'heat exchange' 'between the tube'a'nd 'gas the' upper vfciirier, passing air through the 'fuel reduced the amount of cracking to a'si'iitable degree,
"a'separate liquid or a separate vapor state.
supplied at a can be "blown" out by` exceeding'th'eir designconditions under certain conditions. VConsequently, the vaporizing combustor wouldA appear to be the shortest? and most 'e'tlicient type'of combustor .under-'high altitude"'condiin 'evolving the"`gas Vaporizing combustors' have b'een proposed' previously,
but have been subjecttofvariousdisadvantages. 'One such prior vaporizing'ornbustor 'passed am'ixtu're ofair In vaporizing tube. In the latter construction cracking of the' liquid fuel voccurred as result of lack of 'sufficient pressureto prevent'a"two-statej'condition. '-.The liquid v'portion tends to stay in the'loWerportion of' the heat tube'while the gaseous portion tends 'to 'collect Due to the relatively portion of the tube overheats and 'when droplets of' fuel 'come' in contact therewith, cracking occurs,l resulting in depositsa'ccumulating on the-interior Walls 'of the' tube with "consequent stoppage of the latter. While in the vaporizing tube' with the the Vlarge size of heat exchange apparatus vreiqu'i'red" for the'air-fuel combination in present large combustors' renders heating of both air and fuel Yimpractical.
The present 'invention providesv for placingl 'the liquid fuel under 'a pressure above' its' critical pressure, 'heating the pressurized fuel toabove 'its"criticalfteinpcrature,` and then expanding the heatedifuel through fuel'emiss'ion devices to the combustor. vWhzen'the fuel" s'ata 'tempeiature and pressure above its criticals, it' is in a' corntIirtliis critistate it uniformly fills any enclosure vthat contains it and it alsouniformly Wets the wall surfaces' of such enclosure. Inthis condition'heat'transferfrom"Walls to gris is uniform.
In accordance'with the-present inventionf'liquid fuel is pressure in 'excess of its vcriticalto a heatexchanger associated with a combustion" 'chamber 'Where the heat of combustion-raises th'e temperature "of the fuel toV above its'critical, resulting'inigasiicatio'nof the fuel. Flow-control`mechanisrri`operates during starting of the power plantto preventtlow 'of fuel from 'the' heatexchanger to the' gas" nozzles of'the' combustion chai'nber luntilfthe temperature of 'thefuelin thelheat-exchanger exceeds the critical. n themea'ntime the flow-control mechanism provides for passa-ge of liquid'fuel to other atomizing starting nozzles to initiatecor'nbu'stion.
` Accordingly, lanother object 'of the invention is ,'to
lprovide means for obtaining completely ga'siied fuel with- Yet `another object of the invention is to'provide a 'gasiiied vfuel system including 'gas riiniiin'g-nozzles and "atomizing startiiig nozzles, together yvithf; means "for automatically switching from Pthe starting nozzle s. to the -running nozzles When'the gaseousfuel is availableim las `-will `oe 4apparent vfrom y'the 'following description and "rate' is`reduced. i
claims taken in connection with the accompanying drawings, forming a part of this application, in which:
Fig. l is a side elevational view of a gas turbine power plant for jet propulsion of aircraft and incorporating the present invention;
Fig. 2 is a diagrammatic view of the novel fuel system;
Fig. 3 is a transverse sectional view, taken along the line III-III of Fig. 1, lookng in tlie direction indicated by the arrows;
Fig. 4 is a vaporization chart for MIL-F 5 624;
Fig. 5 is an enlarged fragmentary, sectional view of one of the running nozzles 34 of Fig. 2;
Fig. 6 is an enlarged fragmentary, sectional view of the valve 45;
Fig. 7 is a schematic view of a thermostatic ow control; and,
Fig. 8 is a similar view of a time delay iiow control.
Referring to Fig. 4, there is shown a vaporization chart for MIL- F 5624, as illustrative of a typical fuel for use in aircraft power plants. lf the fuel were at a pressure below its critical pressure of approximately 425 p. s. i. a. and was heated to a temperature sufficient to vaporize it 100 per cent, it would cross all of the curves of the chart. As a result, in passing from the 0 per cent vaporized curve to the 100 per cent vaporized curve the fuel would be in two states, that is, partially liquid and partially gaseous. It is while the fuel is in two states that cracking may occur, resulting in formation of deposits on the walls of the heat-exchanger.
The present invention provides for vaporization of the fuel without passing it through this dangerous twostate stage.
Referring again to the chart of Fig. 4, the two state stage, which exists between the first and last curves of the chart, is avoided by completely by-passing the curves. This is done by first raising the pressure of the fuel to above its critical, that is, to above 425 p. s. i. a. on the chart while its temperature is less than 150 degrees F. This is indicated by the dotted line A. At this point the fuel enters the heat exchanger and its temperature is .increased to above its critical, that is, to above about 670 F., as indicated by the dotted line B. At the right hand end of dotted line B, the fuel is in a critical or single state and is ready to leave the heat exchanger and pass to the fuel nozzles or other fuel-emission devices and hence to the combustion chamber. In so doing its pressure will drop, as indicated by the dotted line C, but at the right of the 100 per cent vaporization curve.
gasifying Referring now to Figs. 1, 2 and 3, for illustration only,
and not by way of limitation, the invention is shown in connection with an aircraft power plant 10 comprising an annular outer casing 11 and an inner coaxial composite core structure 12, the two being radially spaced to provide therebetween an annular passage 13 for iiow of air and gases substantially straight through from an inlet 14 at the front, or left as viewed in Fig. 1, to an exhaust nozzle 15 at the rear or opposite end.
The composite core structure may include a fairing cone 16 at the inlet end, an axial iiow compressor 17, combustion apparatus indicated in its entirety by the reference character 18, a gas turbine 19 and a tail cone 20 which cooperates with the rear end of the outer casing 11 to define the exhaust nozzle 15. The turbine 19 is connected to the compressor 17 and drives the latter through shaft 22 journaled in suitable bearings 23.
A power plant of this character operates in accordance with well known principles, which may be summarized as follows: Air entering the inlet 14 is coinpressed iii the compressor 17 and passes to the combustion apparatus 18 where its temperature is raised by combustion of fuel therein. The heated air and hot products of combustion are expanded through the blading of the turbine 19 to motivate the latter and in turn the compressor 17, the exhaust from the turbine being discharged from the power plant through the exhaust nozzle 15 in the form of a jet for propelling the aircraft in, or on which, the power plant is mounted.
The combustion apparatus, indicated in its entirety by the reference character 18, comprises inner and outer annular walls 26 and 27, respectively, joined at their upstream ends by an annular transverse wall 28. The inner wall 26 is radially spaced fromv a shaft enclosing wall 29, throughout the major portion of the length of the former to deiie therebetween an annular space 31 for flow of compressed air in blanketing relation to the combustion chamber 32 between the inner and outer walls 26 and 27.
Similarly, the outer wall 27 is radially spaced throughout the major portion of its length from the outer casing 11 to define therebetween an airfiow space 33 blanketing the combustion chamber 32 at the outer side of the latter. Preferably, the inner and outer walls 26 and 27 are provided with openings at axially-spaced points for admission of air from the blanketing spaces 31 and 33 to the combustion chamber 32.
Fuel is admitted to the upstream portion of the combustion chamber 32 through an annular series of running nozzles 34 carried by the chamber end wall 28 and supplied with gasified fuel from the common manifold 35.
A fuel pump 36 delivers liquid fuel at a pressure above its critical from a reservoir 37 via conduit 38 to a heat-exchanger 39, preferably disposed in the combustion chamber 32, where the liquid fuel is heated to a temperature above its critical by the heat of combustion occurring within the chamber 32 while at a pressure above its critical. Passage of gasified fuel from the heat exchanger 39 to the manifold 35 is controlled by the flow-control mechanism 41, to be hereinafter more fully described.
For starting the power plant and obtaining sufcient heat to initiate gasification of fuel in the heat-exchanger, a plurality of starting nozzles 42 are provided, these starting nozzles preferably alternating with the running nozzles 34, as best shown in Fig. 3. Liquid fuel is supplied directly to the starting nozzles through manifold 43 b y the pump 36 through a branch 44 of the conduit 38. in bypassing relation to the heat-exchanger.
The flow-control mechanism 41 may involve a thermostatic device (Fig. 7) functioning to retain the valve 45 between the heat-exchanger and the running nozzles in closed position until the temperature of the heat-exchanger exceeds the critical temperature of the fuel therein, when the valve 45 will open and permit flow of gasiiied fuel to the running nozzles. The thermostat 5 1, mounted on the heat exchanger 39, will close the circuit 52 when the desired temperature is reached, thereby operating the solenoid 53 to move the valve 45, through arm 50, to open position, and the valve 46 to corresponding closed position. While the valve 45 in the communication to the running nozzles is retained closed a corresponding valve 46 in the branch conduit 44 t o the starting nozzles will be held open for passage of liquid fuel to the latter nozzles. The flow control mechanism may also include an orifice 45a (Fig. 6) for restricting iiow therethrough to maintain the desired pressure of fuel in the heat-exchanger. However, the fuel nozzles 34 may be provided with conventional orifices 34a for this purpose (Fig. 5).
Instead of relying on a thermostatic device in the fiow-control mechanism, a simple time delay device may be utilized. In such a device, the operating arm of thecontrol valve 45 may be moved to valve opening position by a link or rod 55 connected thereto at one end and carrying at its other end a piston 56 slidable in the chamber 57 of the housing 58. A spring 59 normally maintains the piston 56 in the position shown in Fig. 8. When the fuel pump 46 is started, it directs fuel under pressure through a branch conduit 59 to the space at the right-hand side of the piston 56, thereby moving the latter to the left, as viewed in Fig. 8, against the pressure of the spring 59 until the latter is completely compressed. Inasmuch as the space at the left of the piston 56 is vented to atmosphere through a relatively small bleed port 60, the movement of the piston 56 will consume a period of time which is predetermined by the size of the bleed port 60, this period being so predetermined that upon its completion the temperature of the heat exchanger will be such that the f uel passing thereto and therethrough will be vaporized at a temperature at, or above, the critical temperature of the particular fuel utilized. With such a device, upon positioning of the throttle for starting the valve 45 would be closed and the valve 46 opened. As clearly indicated in Fig. 2, this delayed movement of the valve 45 controls admission of vaporized fuel to the running nozzles 33 will produce a corresponding Closing of the valve 46, thereby shutting ofi flow of liquid fuel to the starting nozzles 42.
As used in this application, the expression conditionresponsive ow control mechanism is intended to include flow control mechanism operated in response to either a thermostatic device or a time delay device, as disclosed above, the two types of devices being substantial equivalents insofar as their use in the present invention is concerned.
While for theoretically perfect functioning of the present invention, critical pressures and temperatures should be utilized, it is recognized that commercially satisfactory results may be obtained at pressures and temperatures as much as to 20 per cent below critical, due primarily to the approaching similarity of conditions of the two states at temperatures and pressures near the critical, and the high turbulence.
While the invention has been shown in but one form, it will be obvious to those skilled in the art that it is not so limited, but is susceptible of various changes and modifications without departing from the spirit thereof.
What is claimed is:
1. In a power plant of the continuous-combustion type, wall structure dening a combustion chamber; one or more running fuel nozzles associated with the combustion chamber for admission of gaseous fuel thereto during operation of the power plant other than starting; one or more starting fuel nozzles associated with the combustion chamber for admission of liquid fuel thereto during starting of the power plant; a source of liquid fuel; pumping mechanism for delivering fuel to the starting nozzle or nozzles and to the running nozzle or nozzles at a pressure near or above the critical pressure of the fuel; a heat-exchanger associated with the combustion chamber and adapted to receive heat from combustion taking place in said combustion chamber; cornmunication means for passage of fuel from the source to the running nozzle or nozzles via the heat-exchanger, whereby the fuel may be heated to a temperature near or above its critical; said communication means also providing for passage of fuel from the source to the starting nozzle or nozzles in bypassing relation to the heat exchanger; and How-control mechanism associated with the communication means operable during starting of the power plant to prevent passage of fuel to the running nozzle or nozzles and to permit passage of fuel to the starting nozzle or nozzles, and operable after starting of the power plant to shut off flow of fuel to the starting nozzle or nozzles and eifect passage of fuel to the running nozzle or nozzles Via the heat-exchanger, whereby the fuel is delivered to the running nozzle or nozzles at a temperature near or above the critical temperature of said fuel.
2. In a power plant of the continuous-combustion type, wall structure defining a combustion chamber and including an annular outer wall; one or more starting nozzles associated with the combustion chamber for admission of fuel thereto during starting of the power plant; one or more running nozzles associated with ,the combustion chamber for admission of fuel thereto after the power plant is operating; a source of fuel; pump mechanism for delivering fuel to the starting nozzle or'nozzles and to the running nozzle or nozzles at the pressure near or above the critical pressure of the fuel; a conduit for passage of fuel from the source to the ruiming nozzle or nozzles, said conduit including a coiled portion associated with the combustion chamber whereby the fuel passing therethrough from the source to the running nozzle or nozzles is heated by combustion occurring in said combustion chamber, said conduit including a branch at a region upstream of the coiled portion, said branch providing for passage of fuel to the starting nozzle or nozzles; condition-responsive flow-control mechanism associated with the conduit at a region downstream of the coiled portion thereof and operable upon starting of the power plant to permit passage of fuel to the starting nozzle or nozzles only via the conduit branch until combustion reaches a degree sufficient to heat fuel in the coiled portion of the conduit to a temperature near or above the critical temperature of the fuel, and thereafter to permit passage of fuel to the running nozzle or nozzles only via the coiled portion of the conduit, and means for maintaining the pressure of the fuel in said coiled portion of the conduit near or above its critical.
References Cited in the file of this patent UNITED STATES PATENTS Number Name Date 1,405,482 Bostedo Feb. 7, 1922 2,479,776 Price Aug. 23, 1949 2,483,045 Harby Sept. 27, 1949 2,502,332 McCollum Mar. 28, 1950 2,506,611 Neal et al. May 9, 1950 2,520,967 Schmitt Sept. 5, 1950 2,591,880 Sammons Apr. 8, 1952
US167115A 1950-06-09 1950-06-09 Liquid fuel vaporizing apparatus Expired - Lifetime US2694899A (en)

Priority Applications (1)

Application Number Priority Date Filing Date Title
US167115A US2694899A (en) 1950-06-09 1950-06-09 Liquid fuel vaporizing apparatus

Applications Claiming Priority (1)

Application Number Priority Date Filing Date Title
US167115A US2694899A (en) 1950-06-09 1950-06-09 Liquid fuel vaporizing apparatus

Publications (1)

Publication Number Publication Date
US2694899A true US2694899A (en) 1954-11-23

Family

ID=22605993

Family Applications (1)

Application Number Title Priority Date Filing Date
US167115A Expired - Lifetime US2694899A (en) 1950-06-09 1950-06-09 Liquid fuel vaporizing apparatus

Country Status (1)

Country Link
US (1) US2694899A (en)

Cited By (16)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
US2916367A (en) * 1955-02-25 1959-12-08 Armstrong Siddeley Motors Ltd Combustion systems for gas turbine engines
US2955420A (en) * 1955-09-12 1960-10-11 Phillips Petroleum Co Jet engine operation
US2958189A (en) * 1955-05-31 1960-11-01 Phillips Petroleum Co Method and apparatus for providing improved combustion in jet engines
US2992527A (en) * 1954-11-17 1961-07-18 Specialties Dev Corp Reaction motor power plant with auxiliary power producing means
US3237401A (en) * 1958-01-17 1966-03-01 United Aircraft Corp Regenerative expander engine
US3365881A (en) * 1965-09-08 1968-01-30 United Aircraft Corp Gas turbine ignition detector
US3367107A (en) * 1965-10-05 1968-02-06 Curtiss Wright Corp Low idle fuel control system
US3447315A (en) * 1967-02-14 1969-06-03 Lucas Industries Ltd Fuel systems for gas turbine engines
DE1601587B1 (en) * 1967-02-16 1972-10-05 Lucas Industries Ltd Fuel supply system for a gas turbine engine
US3707074A (en) * 1970-09-30 1972-12-26 Westinghouse Electric Corp Spontaneous ignition of fuel in a combustion chamber
US3991558A (en) * 1975-12-01 1976-11-16 General Motors Corporation Turbine engine starting fuel control
US5528903A (en) * 1992-03-20 1996-06-25 Schneider-Sanchez Ges.M.B.H. Small gas turbine
US20070180814A1 (en) * 2006-02-03 2007-08-09 General Electric Company Direct liquid fuel injection and ignition for a pulse detonation combustor
US20090151358A1 (en) * 2003-10-23 2009-06-18 United Technologies Corporation Turbine Engine Combustor
US20140238041A1 (en) * 2013-02-27 2014-08-28 General Electric Company Combustor can temperature control system
JP2016502622A (en) * 2012-11-30 2016-01-28 ゼネラル・エレクトリック・カンパニイ Turbine engine assembly and dual fuel aircraft system

Citations (7)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
US1405482A (en) * 1919-05-31 1922-02-07 Louis G Bostedo Method of and means for propelling craft navigating a fluid medium
US2479776A (en) * 1944-04-15 1949-08-23 Lockheed Aircraft Corp Turbo-jet power plant with fuel vaporizer for afterburners
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
US2502332A (en) * 1945-04-12 1950-03-28 Thelma Mccollum Aspirator compressor type jet propulsion apparatus
US2506611A (en) * 1948-03-02 1950-05-09 Westinghouse Electric Corp Fuel control for aviation gas turbine power plants
US2520967A (en) * 1948-01-16 1950-09-05 Heinz E Schmitt Turbojet engine with afterburner and fuel control system therefor
US2591880A (en) * 1948-01-10 1952-04-08 Lockheed Aircraft Corp Liquid and vapor control valve

Patent Citations (7)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
US1405482A (en) * 1919-05-31 1922-02-07 Louis G Bostedo Method of and means for propelling craft navigating a fluid medium
US2479776A (en) * 1944-04-15 1949-08-23 Lockheed Aircraft Corp Turbo-jet power plant with fuel vaporizer for afterburners
US2502332A (en) * 1945-04-12 1950-03-28 Thelma Mccollum Aspirator compressor type jet propulsion apparatus
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
US2591880A (en) * 1948-01-10 1952-04-08 Lockheed Aircraft Corp Liquid and vapor control valve
US2520967A (en) * 1948-01-16 1950-09-05 Heinz E Schmitt Turbojet engine with afterburner and fuel control system therefor
US2506611A (en) * 1948-03-02 1950-05-09 Westinghouse Electric Corp Fuel control for aviation gas turbine power plants

Cited By (21)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
US2992527A (en) * 1954-11-17 1961-07-18 Specialties Dev Corp Reaction motor power plant with auxiliary power producing means
US2916367A (en) * 1955-02-25 1959-12-08 Armstrong Siddeley Motors Ltd Combustion systems for gas turbine engines
US2958189A (en) * 1955-05-31 1960-11-01 Phillips Petroleum Co Method and apparatus for providing improved combustion in jet engines
US2955420A (en) * 1955-09-12 1960-10-11 Phillips Petroleum Co Jet engine operation
US3237401A (en) * 1958-01-17 1966-03-01 United Aircraft Corp Regenerative expander engine
US3365881A (en) * 1965-09-08 1968-01-30 United Aircraft Corp Gas turbine ignition detector
US3367107A (en) * 1965-10-05 1968-02-06 Curtiss Wright Corp Low idle fuel control system
US3447315A (en) * 1967-02-14 1969-06-03 Lucas Industries Ltd Fuel systems for gas turbine engines
DE1601587B1 (en) * 1967-02-16 1972-10-05 Lucas Industries Ltd Fuel supply system for a gas turbine engine
US3707074A (en) * 1970-09-30 1972-12-26 Westinghouse Electric Corp Spontaneous ignition of fuel in a combustion chamber
US3991558A (en) * 1975-12-01 1976-11-16 General Motors Corporation Turbine engine starting fuel control
US5528903A (en) * 1992-03-20 1996-06-25 Schneider-Sanchez Ges.M.B.H. Small gas turbine
US20090151358A1 (en) * 2003-10-23 2009-06-18 United Technologies Corporation Turbine Engine Combustor
US8020366B2 (en) 2003-10-23 2011-09-20 United Technologies Corporation Turbine engine combustor
US8186164B2 (en) 2003-10-23 2012-05-29 United Technologies Corporation Turbine engine fuel injector
EP1526333B1 (en) * 2003-10-23 2013-01-09 United Technologies Corporation Combustor system and fueling method for a gas turbine engine
US20070180814A1 (en) * 2006-02-03 2007-08-09 General Electric Company Direct liquid fuel injection and ignition for a pulse detonation combustor
JP2016502622A (en) * 2012-11-30 2016-01-28 ゼネラル・エレクトリック・カンパニイ Turbine engine assembly and dual fuel aircraft system
US20140238041A1 (en) * 2013-02-27 2014-08-28 General Electric Company Combustor can temperature control system
JP2014163390A (en) * 2013-02-27 2014-09-08 General Electric Co <Ge> Combustor fuel temperature control system
US9303564B2 (en) * 2013-02-27 2016-04-05 General Electric Company Combustor can temperature control system

Similar Documents

Publication Publication Date Title
US2694899A (en) Liquid fuel vaporizing apparatus
US3925002A (en) Air preheating combustion apparatus
US2602289A (en) Method and means for propelling a vehicle using normally gaseous fuel as a liquid
US2552851A (en) Combustion chamber with retrorse baffles for preheating the fuelair mixture
US2390959A (en) Gas turbine power plant
US3002340A (en) Rocket gas generator for turbofan engine
US2501078A (en) Aircraft gas turbine power plant
US2469679A (en) Gas turbine
US3237400A (en) Turborocket engine
US2616254A (en) Jet engine fuel control for modifying fuel pressure drop across throttle in accordance with altitude
US2648950A (en) Gas turbine engine apparatus designed to burn wet pulverized fuel
US2516910A (en) Gas turbine apparatus with selective regenerator control
GB704622A (en) Fuel system for jet and rocket motors
US2446523A (en) Fuel control apparatus for liquid fuel burners
US2712218A (en) Gas turbine apparatus
US2482394A (en) Gas turbine
US2476171A (en) Smoke screen generator
US2755621A (en) Gas turbine installations with output turbine by-pass matching the output turbine pressure drop
US2425630A (en) Internal-combustion airplane heater
US2672727A (en) Fuel vaporizer system for combustion chambers
US3237401A (en) Regenerative expander engine
US3046731A (en) Flame stabilization in jet engines
GB798617A (en) Improvements in aircraft propulsion
US3020709A (en) Control means of the flow of a fluid by another flow
US3348380A (en) Power plant for vtol or stol aircraft having means to augment jet thrust when the same is directed vertically