US3357176A

US3357176A - Twin spool gas turbine engine with axial and centrifugal compressors

Info

Publication number: US3357176A
Application number: US489149A
Authority: US
Inventors: Sam B Williams
Original assignee: Williams International Corp
Current assignee: Williams International Corp
Priority date: 1965-09-22
Filing date: 1965-09-22
Publication date: 1967-12-12
Anticipated expiration: 1984-12-12
Also published as: GB1125251A

Description

1957 s. B. WELLEAMS 3,

TWIN SPOOL GAS TURBINE ENGINE WITH AXIAL AND CBNTRIFUGAL COMPRESSORS Filed Sept. 22, 1965 4 Sheets-Sheet 1 INVENTOR. 1777 ,F, l r z/fiarz'ns OQJMZ Dec. 12, 1967 s. B. WILLIAMS TWIN SPOOL GAS TURBINE ENGINE WITH AXIAL AND CENTRIFUGAL COMPRESSORS 4 Sheets-Sheet 2 Filed Sept. 22, 1965 Dec. 12, 1967 s. B WILLIAMS 3,357,176

TWIN SPOOL GAS TURBINE ENGINE WITH AXIAL 7 AND CENTRIFUGAL COMPRESSORS Filed Sept. 22, 1965 4 Sheets-Sheet 5 1967 s. B. WILLIAMS TWIN SPOOL GAS TURBlNE ENGINE WITH AXIAL AND CENTRIFUG AL COMPRESSORS 4 Sheets-Sheet 4 Filed Sept. 22, 1965 lNv-ENTOR. 64777 E, V/z'Z/zamd United States Patent 3,357,176 TWIN SPGOL GAS TURBINE ENGENE WITH AXIAL AND CENTRIFUGAL COMPRESSORS Sam B. Williams, Walled Lake, Mich, assignor to Williams Research Corporation, Walled Lake, Mich, a corporation of Michigan Filed Sept. 22, 1965, Ser. No. 489,149 2 Claims. (Cl. 6039.16)

ABSTRACT OF THE DISCLOSURE A twin spool gas turbine engine with a low speed axial compressor and a high speed centrifugal compressor, and an annular combustion chamber, with bypass or air bleed features.

This invention relates to gas turbine engines, and more particularly to the compressor sections of such engines.

The invention is concerned with a twin spool arrangement for such compressor sections, in which a first or low pressure spool carries latter stage turbine blades as driving elements and an axial compressor as the driven element, and the second or high pressure spool comprises first stage turbine blades as driving elements and a centrifugal compressor as the driven element, the two spools being coaxially mounted but independently rotatable.

It is an object of the invention to overcome the deficiences inherent in the various previously known compressor arrangements. One of the chief problems of such previous arrangements has been the inability to achieve relatively high pressure ratios (i.e., the ratio of the pressure leaving the compressor to atmospheric pressure) While avoiding surge conditions and inefliciency problems over a broad range of speeds and loads. One such previously known construction utilizes a relatively long axial compressor for both low and high pressures, but it has been found that variable blading is needed in this construction in order to avoid surge at off-design points. Typically, the high pressure ratio single-shaft axial compressor will operate with the early stages stalled at part load when part load is achieved at reduced speed and pressure ratios. Another known compressor arrangement utilizes twin spools for low and high pressures, but with both spools having axial compressors as their driven elements. While avoiding the necessity for variable pitch blades, this arrangement requires extensive miniaturization as Well as fragility and ineiiiciency in the high pressure portion of the compressor, and may create surge problems during acceleration because of the relatively unfavorable pumping characteristics of the high pressure element. It is also necessary in this type of compressor to provide blow-off valves at some intermediate position thereof.

The provision of an axial and centrifugal compressor on one shaft is another conventional arrangement, but this requires the use of a transonic axial compressor, and the axial and centrifugal compressor sections are thismatched at part loads.

It is an object of the present invention to provide a gas turbine engine compressor arrangement which achieves relatively high pressure ratios while avoiding surge problems during acceleration or at off-design points, Without the need for variable pitch blading.

It is another object to provide an improved gas turbine engine compressor arrangement of this character in Which the axial compressor offers the optimum available efficiency for the front portion of the unit and the centrifugal section offers the highest available efliciency (in view of size effects and flow range problems) for the high pressure portion.

It is also an object to provide an improved compressor arrangement having these characteristics, which considerably improves reduced speed part load performance be- Cause the low pressure unit may operate nearer to its optimum or design blade incidences.

It is a further object to provide an improved compressor arrangement which replaces upwards of nine stages of an equivalent axial high pressure spool, and is more efficient than the high pressure axial stage because of their very minute blade size in this application.

It is another object to provide an improved arrangement of this character which is economical to develop and permits elimination of expensive miniaturization and the ability to test each section separately and match all stages easily.

It is also an object to provide an improved gas turbine engine compressor of this nature in which degradation of performance due to erosion of the final stages of an all-axial compressor, or to flow distortion is eliminated by the substitution of the centrifugal compressor, which is more rugged and has been shown to be remarkably insensitive to inlet flow distortion.

It is also an object to provide an improved gas turbine engine compressor of this character which gives the designer the ability to choose the best design for both the axial and centrifugal stages, and permits an optimum speed ratio at varying loads between the high and low pressures, eliminating the need for a transonic axial compressor as when an axial and centrifugal compressor are incorporated on one shaft.

It is another object to provide an improved compressor of this character in which the low pressure section may e operated at a lower speed than the high pressure section, making it possible to devise an optimum second stage turbine design which is not compromised by the requirements for excessive axial driven velocities or of blade root stresses.

It is a further object to provide an improved compressor of this nature which is adaptable to various types of engines and is particularly useful with a concentric or annular combustion chamber, making for a shorter total engine length and thus greater compactness.

Other objects, features and advantages of the present invention will become apparent from the subsequent description, taken in conjunction with the accompanying drawings.

In the drawings:

FIGURE 1 is a longitudinal cross-sectional view of a twin spool fan jet engine embodying the principles of the invention;

FIGURE 2 is a longitudinal cross-sectional view showing a twin spool jet engine with the invention incorporated therein;

FIGURE 3 is a longitudinal cross-sectional view of a twin spool air bleed engine with the compressor arrangement of this invention; and

FIGURE 4 is a longitudinal cross-sectional view of a twin spool gas generator shaft power engine with the principles of the present invention incorporated therein.

Referring more particularly to FIGURE 1 of the drawings, the twin spool fan jet engine is generally indicated at 11 and comprises an annular outer housing 12 and an inner housing 13, the annular space 14 between these housings comprising the air bypass. An air inlet 15 is mounted at the forward end of the engine, and a low pressure axial compressor section generally indicated at 16 is disposed rearwardly of inlet 15. An annular conduit 17 leads from the exit of compressor 16 to a high pressure centrifugal compressor generally indicated at 18. A diffuser 19 leads from the exit of compressor 18 to a space 21 within which is disposed an annular combustion chamber 22. The burned gases leaving chamber 22 lead through nozzle vanes 23 to a first stage turbine 24. The second stage 3 nozzle vanes 25 are disposed on the other side of turbine 24 and lead to second stage turbine 26. The third stage nozzle vanes 27, disposed on the other side of turbine 26, conduct the burned gases to third stage turbine 28. A jet exit 29 is disposed on the other side of turbine 28, inwardly of the bypass exit 31leading from space 14.

The low pressure spool is generally indicated at 32 and comprises a shaft 33 which connects axial compressor 16 with second and third stage turbines 26 and 28 respectively. Compressor 16 comprises four stages the first two stages 34 and 35 having relatively long blades so as to compress the bypass air as well as the air for thecornbustion chamber. All four stages of compressor 16 that is stages 34 and 35 as well as the last two stages 36 and 37, are mounted on a common hub assembly 38 which is secured to a sleeve 39, this sleeve being connected to shaft 33 by a connecting shaft 41. Sets of vanes 42, 43, 44 and 45 are disposedbetween the compressor stages and arev secured to housing portions 12 and 13, the entrance to space 14 having additional vanes 46. Bearings 47 and 48 are provided for rotatably supporting the opposite ends of sleeve 39.

Turbines 26 and 28 have a common hub 49 which is rotatably supported by bearings 51 and 52. This hub is drivably connected at 53 and 54 to shaft 33. V The second or high pressure spool is generally indicated at 55 and comprises tubular shaft 56 which surrounds shaft 33. Bearings 57 and 58 are provided for shaft 56, shaft 33 being disposed inwardly of shaft 56. The hub 59 of centrifugal compressor 18 is secured to shaft 56. Compressor 18 is adapted to change the direction of air flow from axial to radial, the compressed air exiting into diifuser 19 which has vanes 61 and redirects the airt-o an axial fiow into chamber 21. Hub 62 of first stage turbine 24 is also secured to shaft 56. An outer shaft 63. is part of high pressure spool '55, this shaft being spaced outwardly from shaft 56 and extending between turbine 24 and compressor 59. Combustion chamber 22 surrounds. shaft 63, and a member 64 is provided which also surrounds this shaft forwardly of the combustion chamberand has a fuel conduit 65 leading to the inner portionof the combustion chamber.

An accessorydrive 66 may alsobe provided, this drive being connected to the forward end of shaft 56 by gearing 67. m n u In operation, the combustion gases will drive turbines 24, 26 and 2-8 and will then leave through jet exit 29. Turbine 24 will drive high pressure compressor section 18, while turbines 26 and 28 drive low pressure compressor section 16. The latter will also serve to compress the bypass air which leaves through exit 31. I Will thus be observedthat the low, pressure axial compressor section and the high pressure centrifugal compressor section run independently of each other, thus permitting optimum speed ratios at varying loads. Typically, the centrifugal compressor will run proportionately faster at low thrust levels. For exam'ple,.under high thrustconditions therotationalflspeed of. axial compressor section 16 could be perhaps 6 of the centrifugal compressor speed, .whereas, at low thrust power levels it might be only about 20%. I

I FIGURE 2.shows a twin spool jet engine, generally indicated at 101, which isbasically similar tothe fan jet engine of-FIG'URE 1, having a first spool generally indicated at 102 and a second spool generally indicated at 103. Spool 102 comprises a. four-stagelow pressure axial compressor. generally indicated at 104, and second and third stageturbines 105. and106, respectively. Spool 103 comprises a high pressure centrifugal compressor 107 and firststageflurbine 108. As in the previous embodiment, spool .1 02 has a shaft 109 which is concentrically mounted within ashaft 111 of spool 103. i V v I Because the engineshown in FIGURE 2 is not a fan jet epgine, the stages of axial compressor section 104 need not be constructed to compress fan jet air, nor need the housing 112 of the engine be formed to provide a fan jet passage.

The operation of the embodiment in FIGURE 2 will be similar to that of FIGURE 1 as far as the invention is concerned, spools 102 and 103 rotating independently to achieve relatively high pressure ratios while avoiding surge conditions despite load and speed variations.

FIGURE 3 shows a twin spool air bleed engine incorporating the principles of this invention, this engine being indicated generally at 201. The purpose of engine 201 is to provide compressed air for use in industrial applications, such as for power tools on construction projects. The invention comprises a first spool generally indicated at 202 and a second spool generally indicated at 203. Spool 202 comprises second and third stage turbines 204 and 205, respectively, together with a four-stage axial compressor generally indicated at 206, these being connected by a shaft 207. Spool 203 comprises a shaft 208 surrounding shaft 207, one end of this shaft being secured to first-stage turbine 209 and the other end to centrifugal compressor section 211.

Diffuser 212 at the outlet of compressor section 211 is provided with diffuser vanes 213 having apertures 214 extending therethrough. These apertures connect compressed air chamber 215, surrounding combustion chamber 216, with an air bleed conduit 217 mounted on housing 218. An annular connecting chamber 219 within housing 218 receives the compressed air from apertures 214, and conduit 217 is connected to chamber 219 by apertures 221 in housing 218. The engine is provided with a jet exit 222 for the spent combustion gases.

The operation of the twin spools will, in this embodiment of the invention, be similar to that previously described. v v d The compressor arrangement of this invention will coact with the other components of air bleed engine 201 in a manner resulting in important advantages, as cornpared with air bleed engines having a pair of axial compressor sections on twin spools or axial and centrifugal compressor sections on the same spool. More particularly, high pressure spool 203 carrying centrifugal compressor section 211 will tolerate the wide fluctuations in flowrate resulting from variations in the rate of bleed air extraction without encountering compressor surge conditions or major changes in compressor outlet pressure.

FIGURE 4 shows still another engine utilizing the principles of the invention, this being a twin spool gas generator shaft power engine generally indicated at 301. The engine differs from those previously described in thatthe useful power output is in the form of rotational shaft power rather than jet power. The engine is provided with a first or low pressure spool 302 having a four-stage axial compressor 303 driven by second and third stage turbines 304 and 305, respectively. The turbine drives compressor 303 through a shaft 304. Outwardly of this shaft is another shaft 305 forming part of the second or high pressure spool 306. This spool is driven by first stage turbine 307 and drives a centrifugal compressor 308. The connections between the low and high pressure compressors as well as from the outlet of the high pressure compressor to the vicinity of combustion chamber 309 are as described as with respect to the previous embodiments. Engine 301, however, is provided with a fourth stage turbine 311 secured to a shaft 312 which has a rotary power output element 313 secured to the outer end thereof. An exhaust collector 314 is provided for collecting the gases exhausted from turbine 311.

The operation of the FIGURE 4 embodiment, as far as the twin spools are concerned, will be the same as those previously described.

While it will be apparent that the preferred embodi ments of the invention disclosed are well calculated to fulfill the objects above stated, it will be appreciated that the invention is susceptible to modification, variation and change without departing from the proper scope or fair meaning of the subjoined claims.

What is claimed is:

1. In a gas turbine engine, a multistage low pressure axial compressor, a high pressure centrifugal compressor, a connecting passage between said compressors, a diffuser extending outwardly from the exit of said centrifugal compressor, an annular combustion chamber rearwardly of said pressure compressor, the outer diameter of said combustion chamber being less than the diameter of said diffuser, a compressed air chamber surrounding said combustion chamber, said difiFuser having apertured vanes connecting said compressed air chamber with an air bleed chamber, an air bleed connection leading outwardly from said air bleed chamber, said diffuser having an axially directed exit leading to said compressed air chamber, a first-stage turbine, a shaft connecting said firststage turbine to said centrifugal compressor, the inner diameter of said combustion chamber being closely adjacent said shaft, the size of said combustion chamber in an axial direction being substantially less than the distance between its inner and outer diameters, a fuel conduit leading to the inner portion of said combustion chamber, a latter-stage turbine, and a shaft connecting said last mentioned turbine with said axial compressor, said second shaft being within and rotatable independently of said first shaft.

2. In a fan jet engine, a multi-stage low pressure axial compressor, a latter stage turbine, a connecting shaft between said latter stage turbine and axial compressor, an annular conduit leading from said axial compressor, an annular bypass air passage outwardly of said conduit, at least one stage of said axial compressor having relatively long blades supplying air to said bypass air passage as well as to said last-mentioned conduit, a high pressure centrifugal compressor connected to said conduit, a firststage turbine, a second shaft connecting said first-stage turbine and the centrifugal compressor and surrounding said first-mentioned shaft, the two shafts being independently rotatable, a diffuser extending outwardly from the exit of said centrifugal compressor, an annular combustion chamber rearwardly of said high pressure compressor, the outer diameter of said combustion chamber being less than the outer diameter of said diffuser, said diffuser having an axially directed exit leading to an annular air space which surrounds said chamber, the inner diameter of said combustion chamber being closely adjacent said second shaft, the size of said combustion chamber in an axial direction being substantially less than the distance between its inner and outer diameters, and a fuel conduit leading to the inner portion of said combustion chamber.

References Cited UNITED STATES PATENTS 2,746,246 5/1956 Valota -39-16 2,952,973 9/1960 Hall et al 60-39.16 X 2,969,644 1/1961 Williams et al. 6039.1 6 3,093,969 6/1963 Moellmann 6039.16 X 3,232,053 2/1966 Rogers et al. 60-39.l6 X

FOREIGN PATENTS 594,139 11/ 1947 Great Britain. 697,285 9/ 1953 Great Britain.

JULIUS E. WEST, Primary Examiner.

Claims

1. IN A GAS TURBINE ENGINE, A MULTISTAGE LOW PRESSURE AXIAL COMPRESSOR, A HIGH PRESSURE CENTRIFUGAL COMPRESSOR, A CONNECTING PASSAGE BETWEEN SAID COMPRESSORS, A DIFFUSER EXTENDING OUTWARDLY FROM THE EXIT OF SAID CENTRIFUGAL COMPRESSOR, AN ANNULAR COMBUSTION CHAMBER REARWARDLY OF SAID PRESSURE COMPRESSOR, THE OUTER DIAMETER OF SAID COMBUSTION CHAMBER BEING LESS THAN THE DIAMETER OF SAID DIFFUSER, A COMPRESSED AIR CHAMBER SURROUNDING SAID COMBUSTION CHAMBER, SAID DIFFUSER HAVING APERTURED VANES CONNECTING SAID COMPRESSED AIR CHAMBER WITH AN AIR BLEED CHAMBER, AN AIR BLEED CONNECTION LEADING OUTWARDLY FROM SAID AIR BLEED CHAMBER, SAID DIFFUSER HAVING AN AXIALLY DIRECTED EXIT LEADING TO SAID COMPRESSED AIR CHAMBER, A FIRST-STAGE TURBINE, A SHAFT CONNECTING SAID FIRSTSTAGE TURBINE TO SAID CENTRIFUGAL COMPRESSOR, THE INNER DIAMETER OF SAID COMBUSTION CHAMBER BEING CLOSELY ADJACENT SAID SHAFT, THE SIZE OF SAID COMBUSTION CHAMBER IN AN AXIAL DIRECTION BEING SUBSTANTIALLY LESS THAN THE DISTANCE BETWEEN ITS INNER AND OUTER DIAMETERS A FUEL CONDUIT LEADING TO THE INNER PORTION OF SAID COMBUSTION CHAMBER, A LATTER-STAGE TURBINE, AND A SHAFT CONNECTING SAID LAST MENTIONED TURBINE WITH SAID AXIAL COMPRESSOR, SAID SECOND SHAFT BEING WITHIN AND ROTATABLE INDEPENDENTLY OF SAID FIRST SHAFT.