Gas turbine engine and the starting thereof

Gas turbine engine and the starting thereof.

Description
December 14, 1954 F. c . MARCHANT RRM . 2GAS TURBINE ENGINE AND STARTING THE SAME Filed August 13, 1948 4 Sheets sheets l 5:00 a.m. mm ML N nw 5 F. c . l . MARCHANT ETAL 2696711GAS TURBINE ENGINE AND THE SAME Filed August 13, 1948 4 Sheets sheets 3 14 December I954 F. c . l . MARCHANT ETAL 2696711GAS TURBINE ENGINE AND STARTING THE SAME Filed August 13, 1948 4 Sheets - Page 4 / / Yl f/V7 '015 E63 1. [ Mew/4md one. 44 . 45mm U.S. Patent GAS TURBINE ENGINE AND SAME STARTING Ivor Francis Charles Manns Marchant and Lewis Gordon , Bristol , England , assignors to the Bristol Aeroplane Company Limited , Bristol , England , a British company application 13 of August 1948 , serial number 44014
2 Claims . ( Cl 60-3914 )
 This invention relates to gas turbine engines and is intended to improve the qualities of starting and operating the engine.According to this invention, a gas turbine engine having a pair of turbines through which the working fluid passes in succession and compressor assembly composed of a unit which is independently driven by a turbine , characterized in that there is provided a starter motor for driving said compressor unit only and all or some of the engine accessories .The present invention finds particular application to gas turbine engines for driving airplanes and according to another feature of the invention , an engine of this type having a low compressor and high pressure in each of which is independently powered by a turbine through which the working fluid passes in succession characterized in that a starter motor for driving only the high pressure compressor and turbine and engine accessories which are driven with the high pressure compressor and the accessories are driven ticket with the low pressure compressor .A practical application of the present invention will now be described , by way of example with reference to the accompanying drawings , which :Figure 1 is a schematic sectional elevation of a gas turbine engine according to the present invention,Figure 2 is a view on an enlarged scale of part of the engine of Figure 1 ,Figure 3 is a cross section showing a part of the motor of Figure l and an enlarged scale, andFigure 4 is an elevational view in section showing a detail construction of the output of the compressor assembly .
Referring to Figure l:

The engine comprises a pair of turbines 5 and 6 which are part of the turbine generally indicated at 7. The engine also comprises a compressor assembly indicated generally at 8, and the combustion apparatus , indicated generally at 9, which connects the compressor assembly 8 with the turbine assembly 7 .The combustion apparatus comprises an annular combustion chamber 10 which contains a plurality of flame tubes 11, the fuel is supplied individually to each of the flame tube in a known manner , the fuel burns in the flame tube and the products of combustion passing along the annular combustor 10 to the turbine assembly 7 . The combustion air is obtained from the assembly of compressor 8 . Combustion products pass successively through the turbines 5 and 6 , respectively.Compressor assembly 8 comprises a low-pressure axial flow compressor 12 and a high pressure axial flow compressor 13. The compressor 12 supplies the compressor 13 which in turn supplies the combustion device 9.The air from the compressor 12 is transmitted to compressor 13 through an annular duct 14 which lies between the compressors. Stator housings of the compressors 12 and 13 are bolted to the duct 14. The compressor 13 is connected to the combustion apparatus by a housing 9 annular air duct 15.Be seen in Figure 1 that the housing 15 is bolted to the stator housing 13 and the compressor outer wall of the annular combustor 10 . The arrangement described thus provides continuous outer layer that surrounds the motor of the compressor outlet to the outlet end comice turbine - the ring housing 16 of the turbine forming part of the shell.The diameter of the shell decreases from the compressor inlet to the annular combustion chamber , and then increased so that the motor shell has a waist portion formed by the air duct housing 15.The high pressure compressor 13 is coupled to the high pressure turbine 5 for a drive shaft 17 and , similarly, the low pressure compressor 12 is connected to the low pressure turbine 6 by a drive shaft 18 , this latter is nested within the hollow shaft 17.The engine is being described is particularly adapted for use in an aircraft and , as is well known , and the plane theengine each requires certain accessories for its operation . For example, the engine requires a starter , lubrication pumps , fuel pumps , and so on while the aircraft requires vacuum pumps , generators, pumps, pressure and so on.An engine constructed in accordance with this invention, the engine starter motor and some other access - ' accessory are carried by the housing 15 and are coupled through suitable gearing with the axle 17. This gear is indicated schematically in Figure l at 19.The detailed construction of the gear 19 of Figure l IINS shown in Figures 2 and 3 to which reference will now e had.The shaft 17 carries a pinion 20 which meshes with a gear wheel 21 mounted on an intermediate shaft 22. The latter is supported by bearings on a frame 23 which is integrally formed with a diaphragm 24 that is bolted , at 25 , to the inner wall of the housing 15. The intermediate shaft 22 carries a bevel gear 26 meshing with a bevel pinion 27 carried by a shaft 28 . The latter is the drive shaft of the starter motor 29 which is bolted to the top of the housing 15 . Thereby the starter motor is connected to the axle 17.A similar gear train 20, 30 , 31, 32 is provided for connecting the shaft 17 with a drive shaft 33 extending down to the bottom of the housing 15 .Bolted to the underside of the housing 15 is a gear box 34 which carries a pair of fuel pumps 35, the pressure pump 36 and a recovery pump 37 . Sweeping and pressure pumps 36, 37 are driven directly by the shaft 33 . Fuel pumps 35 are driven through a gear train 38, 39 , 40, 41 carried within the gearbox 34 .Therefore, it is seen that the starter motor 29 and some accessories 35 , 36 and 37 are coupled to shaft 17 by a gear train (located within the inner wall of the housing 15 ) and the axles radially arranged 28, 33 . Inner and outer walls of the housing 15 are interconnected by a series of hollow radial arms 42 and each of the shafts 28 , 33 passes through one of said arms . The arms 42 are streamlined in the direction of air flow from the compressor to the combustion chamber .With the described arrangement , when it is desired to start the engine starter motor 29 is operated to drive the shaft 17. As a result , the compressor temperature HIGH - pressure turbine 13 and 5 may be rotated while the engine accessories 35 , 36 and 37 are actuated . These accessories include startlng and essential for engine operation. When starting the combustion of high pressure turbine 5 starts driving the compressor 13 and engine accessories . Thus, the starting operation takes gently on the impeller 5 : At the same time, the turbine 6 begins to drive the low pressure compressor 12.Incorporated in the coupling between the starter motor 29 and the shaft 17 is a known construction of free-wheel device to ensure that when the impeller 5 is set in motion the shaft 17 is free to run over the starter motor 29.It is preferred that aircraft accessories not to be taken to start the engine as this requires a larger starter motor horsepower .Accordingly, aircraft accessories are mounted on a gear box supported by a portion of the aircraft convenient and driven by a radial shaft connected to low pressure shaft 18. Such arrangement is shown schematically in Figure 1 in which the aircraft accessories 43, 44, 45 and 46 are carried by a gearbox 47 which is driven by a shaft 48 . The latter is coupled to shaft 18 by a gear train generally indicated at 49 -, - the shaft 48 passes through a hollow radial arm 50 which connects the inner and outer walls of the duct 14.When the turbine i begins. turbinewill operateunderits own power to low pressure is . implemented to boost compressor'll as indicated above . When this happens the whole enginelas power delivery is so wide . J is then available. to conduct . aircraft accessories 43-47 . iii The arms 42, 50 that extendacross walls of the housing and direct the conduit '14 respectively . airaxially through the engine , it is necessary that the circumferential component cira theair ofmovementof out . the compressor 13 and reduce orfeliminated . to .. this "reason of . output . of theicompressor . 13 is provided with two rows of stator. Vanes51 , 52 ( see - Eigurefi ) lying one immediately behind the other . La . statorvanes 51 reduce the circumferential component and direct air . really along an axial path wheretheprocedure in'tofrow 52 repeats. Leaving the air. Thecompressoriassembly 8 engine flows substantially longitudinally .Dos. stator rows . vanes 51, 5.2tare provided to straighten the flow of air from the circumferential component to be eliminatedislarge . Provisionofztwo th blade rows in such aLconditiondoes . The'compressor adiversely Afiect not efficien'cy'as whenansingle straightening blade row is provided . 
The provision of two rows. Stator . vanes has the advantage that the two rows of blades can be relatively adjusted so that to ensure the most efficient flowiof ainthrough palettes. For example, it may be more advantageousto . the downstream rowofvanes space . theupstream circumferential row of vanesso vanes52 found between the vanes 51 centrallyor the contrary. With this arrangement also higher. spacing . such vanes and allowed for . the combined effect is . - That the air flow is turned in the axial direction with less loss than if provided asingle row of blades ;It is understood that the invention is not'restricted a gas turbine engine compressor having low axial flow and'highpressure . -8 Theasse'rnb1y may consist of a axial -flow compressor in the low pressure delivery of a high pressure centrifugal compressor , or alternatively, may consist of a low pressure compressor axial flow gives andcentrifugalcornpressor acombinedaxial .
We claim :

1. A gas turbine engine for the propulsion of an aircraft comprising a multi-stage turbine , a compressor stage multistage each of which is independently powered by a turbine rotor , compressor and turbine that has a common rotation axis and spaced longitudinally thereof, combustion equipment in the space between the compressor and air turbine cornpressor tojreceive and to deliver products of combustion to the turbine, accessories needed for engine operation operation, it is driven by the high pressure compressor , accessories necessary for operation of the aircraft , said accessories being driven by the rotor of the turbine to a low pressure compressor , a starter motor to drive said compressor only high pressure and engine accessories , yun shell that surrounds the engine provided with a waist portion having an outer diameter smaller than the diameter over- all was said , shell , said said starter motor and engine accessories are arranged around waist , he said. portioinof shell that is positioned substantially in diameter on - all out of the shell.Two . Argas turbineengine as claimed in claim 1 characterized by, that said comprises a housing waistportion air duct connecting the high pressure compressor to said combustion device , said housing has concentric inner and outerwalls and radialarms extends through the walls, atleasttwo ofsaid armsbeing ' each recess for receiving a shaft connecting said starter motor and said engineflaccessories the drive shaft between the high pressure - turbineand its compressor.