US3385065A - Gas turbine jet propulsion engine - Google Patents

Gas turbine jet propulsion engine Download PDF

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US3385065A
US3385065A US632828A US63282867A US3385065A US 3385065 A US3385065 A US 3385065A US 632828 A US632828 A US 632828A US 63282867 A US63282867 A US 63282867A US 3385065 A US3385065 A US 3385065A
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compressor
gas turbine
nozzles
axial
compressor means
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US632828A
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Coplin John Frederick
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Rolls Royce PLC
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Rolls Royce PLC
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    • F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F02—COMBUSTION ENGINES; HOT-GAS OR COMBUSTION-PRODUCT ENGINE PLANTS
    • F02K—JET-PROPULSION PLANTS
    • F02K3/00—Plants including a gas turbine driving a compressor or a ducted fan
    • F02K3/02—Plants including a gas turbine driving a compressor or a ducted fan in which part of the working fluid by-passes the turbine and combustion chamber
    • F02K3/025—Plants including a gas turbine driving a compressor or a ducted fan in which part of the working fluid by-passes the turbine and combustion chamber the by-pass flow being at least partly used to create an independent thrust component
    • B—PERFORMING OPERATIONS; TRANSPORTING
    • B64—AIRCRAFT; AVIATION; COSMONAUTICS
    • B64C—AEROPLANES; HELICOPTERS
    • B64C29/00—Aircraft capable of landing or taking-off vertically, e.g. vertical take-off and landing [VTOL] aircraft
    • B64C29/0008—Aircraft capable of landing or taking-off vertically, e.g. vertical take-off and landing [VTOL] aircraft having its flight directional axis horizontal when grounded
    • B64C29/0041—Aircraft capable of landing or taking-off vertically, e.g. vertical take-off and landing [VTOL] aircraft having its flight directional axis horizontal when grounded the lift during taking-off being created by jet motors
    • B64C29/0066—Aircraft capable of landing or taking-off vertically, e.g. vertical take-off and landing [VTOL] aircraft having its flight directional axis horizontal when grounded the lift during taking-off being created by jet motors with horizontal jet and jet deflector
    • F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F02—COMBUSTION ENGINES; HOT-GAS OR COMBUSTION-PRODUCT ENGINE PLANTS
    • F02C—GAS-TURBINE PLANTS; AIR INTAKES FOR JET-PROPULSION PLANTS; CONTROLLING FUEL SUPPLY IN AIR-BREATHING JET-PROPULSION PLANTS
    • F02C7/00—Features, 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
    • F02C7/04—Air intakes for gas-turbine plants or jet-propulsion plants

Definitions

  • the engine being provided with a completely separate and distinct axial compressor means for supplying air to vectorable nozzles mounted forwardly of the normal propulsion nozzles.
  • the distinct and separate compressor means for the vectorable nozzles is driven by the compressor means for the forward propulsion nozzles, the designs of each compressor means being the most optimum for its particular function.
  • This invention concerns a gas turbine engine.
  • a gas turbine engine comprising a first casing which defines an inner duct; first axial flow compressor means, combustion equipment, and turbine means arranged in flow series in said inner duct; a second casing which is arranged concentrically around the first casing and which defines therewith an outer duct; second axial fiow compressor means arranged in said outer duct; and first and second propulsion nozzles arranged at the downstream ends of the inner and outer ducts respectively; the second axial flow compressor means having a plurality of stages of rotor blades each of which stages is drivingly connected to or integral with corresponding stages of the first axial flow compressor means.
  • both the first axial flow compressor means and of the second axial fiow compressor means will, in practice, be small, the length of both the first and second axial flow compressor means will be reduced for any given blade aspect ratio.
  • the second axial flow compressor means will have a higher root blade speed than the first axial fiow compressor means, the second axial flow compressor means will require fewer stages than the first axial fiow compressor means to give the same overall pressure ratio.
  • the two second propulsion nozzles may, for example, be aligned with the said combustion equipment.
  • the second propulsion nozzles may, without the use of volutes or reverse flow ducts, be arranged further forward than would otherwise be the case, and such forward disposition of these nozzles is desirable in order to effect a balance, about the centre of gravity of the aircraft, of the thrusts produced by said nozzles.
  • the outer duct is provided with a diffusing portion which is disposed between the second axial fiow compressor means and the said second propulsion nozzles.
  • Each of the said second propulsion nozzles is preferably adjustable to vary the direction of flow, of the air 3,335,955 Patented May 28, 1968 leaving the nozzle, progressively between vertical and horizontal directions.
  • the axial velocity of the air leaving the downstream end of the second axial fiow compressor means may be less than that of the air leaving the downstream end of the first axial flow compressor means.
  • Means may be provided for sealing the inner and outer ducts from each other.
  • the said turbine means comprises a high pressure turbine and a low pressure turbine, the high pressure turbine driving a high pressure compressor, and the low pressure turbine driving a low pessure compressor which constitutes the said first axial flow compressor means.
  • FIGURE 1 is a part sectional diagrammatic view of a gas turbine jet propulsion engine according to the present invention.
  • FIGURE 2 is a broken-away sectional view of part of the structure shown in FIGURE 1.
  • a gas turbine engine 16 is provided with a casing 11 which defines an inner duct 12 which has, at its upstream end, an air intake 13.
  • a low pressure compressor 14 mounted in fiow series in the inner duct 12 are a low pressure compressor 14, a high pressure compressor 15, combustion equipment 16, a high pressure turbine 17, and a low pressure turbine 18.
  • the turbine exhaust gases are directed to atmosphere through a bifurcated exhaust duct 19 which constitute the downstream portion of the inner duct 12.
  • the exhaust duct 19 is provided with two diametrically spaced apart propulsion nozzles 20.
  • the high pressure turbine 17 drives the high pressure compressor 15 through a shaft 21.
  • the low pressure turbine 18 drives the low pressure compressor 14 through a shaft 22 which is mounted concentrically within the shaft 21.
  • a casing 23 is arranged concentrically around the easing 11 and defines therewith an outer duct 24 at whose upstream end there is an air intake 25.
  • the low pressure compressor 14 is provided with four stages 26 of rotor blades, of which the first three stages only are formed integrally with stages 27 of rotor blades of a compressor 28 arranged in the outer duct 24.
  • labyrinth seals 30 are provided between the casing 11 and the junctions of the blades 25, 27.
  • the axial velocity of the air leaving the downstream end of the compressor 28 is less than that of the air leaving the downstream end of the compressor 14 whereby the compressor 28 has a higher tip speed for a given rotor incident mach number.
  • the outer duct 24 has a diffusing portion 31 which is disposed immediately downstream of the compressor 23.
  • the diffusing portion 31 leads to two diametrically spaced apart propulsion nozzles 32 which are aligned with the combustion equipment 16 and which are thus arranged substantially forwardly of the propulsion nozzles 20.
  • Each of the propulsion nozzles 32 may be rotated (by means not shown) so as to vary the direction of flow of the air leaving these nozzles progressively between vertical and horizontal directions.
  • the construction of the gas turbine engine shown in the drawings is such that, by reason of the forward disposition of the nozzles 32, it may be mounted in an aircraft in a position in which the thrusts from the nozzles 20, 32 are substantially balanced about the centre of gravity of the aircraft. Prior to the present invention, however, it has been necessary, in order to position the nozzles 32 forwardly, to employ either a volute or a reverse flow duct. As will be seen, the construction shown in the drawings does not involve the use of such volutes or reverse flow ducts, and is lighter and more compact in consequence.
  • the ability to dispense with such volutes and reverse flow ducts arises from the short axial length of the compressor 28 and this in turn arises from the fact that the small radial extent of the compressors 14, 23 enables the length of both these compressors to be reduced for any given blade aspect ratio.
  • the axial length of the compressor 28 is especially short by reason of the fact that the compressor 23 is working at a higher root blade speed than the compressor 14, With the result that it can achieve the same pressure ratio in a reduced number of stages. Greater length is therefore provided for diffusion.
  • the engine is particularly well balanced about its centre of lift.
  • a gas turbine engine comprising: a first casing which defines an inner duct; a rotatable first low pressure axial flow compressor means having a plurality of stages of rotor blades, a high pressure compressor means rotatable independently of said first compressor means, combustion equipment, a high pressure turbine means, and a low pressure turbine means arranged in flow series in said inner duct; said high pressure turbine means driving said high pressure compressor means and said low pressure turbine means driving said plurality of stages of rotor blades of said low pressure compressor means; a first propulsion nozzle arranged at the downstream end of the inner duct; a second casing arranged concentrically around the first casing and defining therewith an outer duct; means for sealing the inner and outer ducts from each other, a second low pressure axial fiow compressor means arranged in said outer duct and having a plurality of stages of rotor blades, each of said stages of rotor blades of said second low pressure compressor means being respectively secured to corresponding stages of rotor blades of said first low pressure
  • a gas turbine engine as claimed in claim 1 in which the two second propulsion nozzles are diametrically spaced apart on the second casing and are located substantially forwardly of the said first propulsion nozzle.
  • a gas turbine engine as claimed in claim 2 in which the two second propulsion nozzles are aligned with the said combustion equipment transversely of the longitudinal axis of the engine.
  • a gas turbine engine as claimed in claim 1 in which the axial velocity of the air leaving the downstream end of the second axial fiow compressor means is less than that of the air leaving the downstream end of the first axial flow compressor means.

Description

y 1968 J. F. COPLIN 3,385,065
GAS TURBINE JET PROPULSION ENGINE Original Filed Jan. 22, 1965 Inventor United States Patent 3,385,065 GAS TURBINE JET PROPULSION ENGINE John Frederick Coplin, Littieover, England, assignor to Rolls-Royce Limited, Derby, Derbyshire, England, a company of Great Britain Continuation of application Ser. No. 427,302, Jan. 22, 1965. This application Apr. 21, 1967, Ser. No. 632,828 Claims priority, application Great Britain, Feb. 21, 1964, 7,468/ 64 5 Claims. (Cl. 60226) ABSTRACT OF THE DISCLOSURE A gas turbine engine of the vectorable thrust type used in V.T.O.L. and S.T.O.L. aircraft, the engine being provided with a completely separate and distinct axial compressor means for supplying air to vectorable nozzles mounted forwardly of the normal propulsion nozzles. The distinct and separate compressor means for the vectorable nozzles is driven by the compressor means for the forward propulsion nozzles, the designs of each compressor means being the most optimum for its particular function.
This application is a continuation of my United States application Ser. No. 427,302, filed Jan. 22, 1965 (now abandoned).
This invention concerns a gas turbine engine.
According to the present invention, there is provided a gas turbine engine comprising a first casing which defines an inner duct; first axial flow compressor means, combustion equipment, and turbine means arranged in flow series in said inner duct; a second casing which is arranged concentrically around the first casing and which defines therewith an outer duct; second axial fiow compressor means arranged in said outer duct; and first and second propulsion nozzles arranged at the downstream ends of the inner and outer ducts respectively; the second axial flow compressor means having a plurality of stages of rotor blades each of which stages is drivingly connected to or integral with corresponding stages of the first axial flow compressor means.
Since the radial dimension of both the first axial flow compressor means and of the second axial fiow compressor means, will, in practice, be small, the length of both the first and second axial flow compressor means will be reduced for any given blade aspect ratio. Moreover, since the second axial flow compressor means will have a higher root blade speed than the first axial fiow compressor means, the second axial flow compressor means will require fewer stages than the first axial fiow compressor means to give the same overall pressure ratio.
Preferably there are two diametrically spaced apart second propulsion nozzles which are arranged substantially forwardly of the said first propulsion nozzle or nozzles. Thus, the two second propulsion nozzles may, for example, be aligned with the said combustion equipment.
Since the length of the second axial flow compressor means will be reduced, as explained above, the second propulsion nozzles may, without the use of volutes or reverse flow ducts, be arranged further forward than would otherwise be the case, and such forward disposition of these nozzles is desirable in order to effect a balance, about the centre of gravity of the aircraft, of the thrusts produced by said nozzles.
Preferably the outer duct is provided with a diffusing portion which is disposed between the second axial fiow compressor means and the said second propulsion nozzles.
Each of the said second propulsion nozzles is preferably adjustable to vary the direction of flow, of the air 3,335,955 Patented May 28, 1968 leaving the nozzle, progressively between vertical and horizontal directions.
The axial velocity of the air leaving the downstream end of the second axial fiow compressor means may be less than that of the air leaving the downstream end of the first axial flow compressor means.
Means may be provided for sealing the inner and outer ducts from each other.
Preferably, the said turbine means comprises a high pressure turbine and a low pressure turbine, the high pressure turbine driving a high pressure compressor, and the low pressure turbine driving a low pessure compressor which constitutes the said first axial flow compressor means.
The invention is illustrated, merely by way of example, in the accompanying drawings, in which:
FIGURE 1 is a part sectional diagrammatic view of a gas turbine jet propulsion engine according to the present invention, and
FIGURE 2 is a broken-away sectional view of part of the structure shown in FIGURE 1.
Referring to the drawings, a gas turbine engine 16 is provided with a casing 11 which defines an inner duct 12 which has, at its upstream end, an air intake 13. Mounted in fiow series in the inner duct 12 are a low pressure compressor 14, a high pressure compressor 15, combustion equipment 16, a high pressure turbine 17, and a low pressure turbine 18. The turbine exhaust gases are directed to atmosphere through a bifurcated exhaust duct 19 which constitute the downstream portion of the inner duct 12. The exhaust duct 19 is provided with two diametrically spaced apart propulsion nozzles 20.
The high pressure turbine 17 drives the high pressure compressor 15 through a shaft 21. The low pressure turbine 18 drives the low pressure compressor 14 through a shaft 22 which is mounted concentrically within the shaft 21.
A casing 23 is arranged concentrically around the easing 11 and defines therewith an outer duct 24 at whose upstream end there is an air intake 25.
The low pressure compressor 14 is provided with four stages 26 of rotor blades, of which the first three stages only are formed integrally with stages 27 of rotor blades of a compressor 28 arranged in the outer duct 24.
In order to prevent flow from the inner duct 12 to the outer duct 24, labyrinth seals 30 are provided between the casing 11 and the junctions of the blades 25, 27.
The axial velocity of the air leaving the downstream end of the compressor 28 is less than that of the air leaving the downstream end of the compressor 14 whereby the compressor 28 has a higher tip speed for a given rotor incident mach number.
The outer duct 24 has a diffusing portion 31 which is disposed immediately downstream of the compressor 23. The diffusing portion 31 leads to two diametrically spaced apart propulsion nozzles 32 which are aligned with the combustion equipment 16 and which are thus arranged substantially forwardly of the propulsion nozzles 20. Each of the propulsion nozzles 32 may be rotated (by means not shown) so as to vary the direction of flow of the air leaving these nozzles progressively between vertical and horizontal directions.
The construction of the gas turbine engine shown in the drawings is such that, by reason of the forward disposition of the nozzles 32, it may be mounted in an aircraft in a position in which the thrusts from the nozzles 20, 32 are substantially balanced about the centre of gravity of the aircraft. Prior to the present invention, however, it has been necessary, in order to position the nozzles 32 forwardly, to employ either a volute or a reverse flow duct. As will be seen, the construction shown in the drawings does not involve the use of such volutes or reverse flow ducts, and is lighter and more compact in consequence.
The ability to dispense with such volutes and reverse flow ducts arises from the short axial length of the compressor 28 and this in turn arises from the fact that the small radial extent of the compressors 14, 23 enables the length of both these compressors to be reduced for any given blade aspect ratio. Moreover, the axial length of the compressor 28 is especially short by reason of the fact that the compressor 23 is working at a higher root blade speed than the compressor 14, With the result that it can achieve the same pressure ratio in a reduced number of stages. Greater length is therefore provided for diffusion.
Moreover, by reason of the short axial length of the compressors 14, 28 and by reason of the disposition of the nozzles 20, 32, the engine is particularly well balanced about its centre of lift.
What is claimed is:
1. A gas turbine engine comprising: a first casing which defines an inner duct; a rotatable first low pressure axial flow compressor means having a plurality of stages of rotor blades, a high pressure compressor means rotatable independently of said first compressor means, combustion equipment, a high pressure turbine means, and a low pressure turbine means arranged in flow series in said inner duct; said high pressure turbine means driving said high pressure compressor means and said low pressure turbine means driving said plurality of stages of rotor blades of said low pressure compressor means; a first propulsion nozzle arranged at the downstream end of the inner duct; a second casing arranged concentrically around the first casing and defining therewith an outer duct; means for sealing the inner and outer ducts from each other, a second low pressure axial fiow compressor means arranged in said outer duct and having a plurality of stages of rotor blades, each of said stages of rotor blades of said second low pressure compressor means being respectively secured to corresponding stages of rotor blades of said first low pressure axial flow compressor means, said second low pressure axial flow compressor means having a shorter axial length than the axial length of said first low pressure axial fiow compressor means; a diffuser portion in said outer duct downstream of the second compressor means; two second propulsion nozzles arranged at the downstream end of said diffuser portion; and means for adjusting said second propulsion nozzles to vary the direction of flow of the air leaving the nozzles progressively between vertical and horizontal directions.
2. A gas turbine engine as claimed in claim 1 in which the two second propulsion nozzles are diametrically spaced apart on the second casing and are located substantially forwardly of the said first propulsion nozzle.
3. A gas turbine engine as claimed in claim 2 in which the two second propulsion nozzles are aligned with the said combustion equipment transversely of the longitudinal axis of the engine.
4. A gas turbine engine as claimed in claim 1 in which the second axial flow compressor means has fewer stages of rotor blades than the first axial flow compressor means.
5. A gas turbine engine as claimed in claim 1 in which the axial velocity of the air leaving the downstream end of the second axial fiow compressor means is less than that of the air leaving the downstream end of the first axial flow compressor means.
References Cited UNITED STATES PATENTS 2,509,890 5/1950 Stalker 230122 3,153,906 10/1964 Marchant 226 3,182,898 5/1965 Hewson 230122 3,191,886 6/1965 Lewis et a1. 60-226 FOREIGN PATENTS 487,628 12/1953 Italy.
MARTIN P. SCHWADRON, Primary Examiner.
D. HART, Assistant Examiner.
US632828A 1964-02-21 1967-04-21 Gas turbine jet propulsion engine Expired - Lifetime US3385065A (en)

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Cited By (2)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
US3545209A (en) * 1967-06-19 1970-12-08 Rolls Royce Gas turbine by-pass engine
US3867813A (en) * 1971-10-05 1975-02-25 Mtu Muenchen Gmbh Turbojet engine for vertical or short take-off and landing airplanes

Families Citing this family (1)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
FR2695960B1 (en) * 1992-09-23 1994-10-21 Snecma Variable cycle engine for a vertical takeoff aircraft.

Citations (4)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
US2509890A (en) * 1945-03-13 1950-05-30 Edward A Stalker Jet-propelled aircraft with boundary layer control
US3153906A (en) * 1960-12-15 1964-10-27 Bristol Siddeley Engines Ltd Aircraft propulsion power units
US3182898A (en) * 1962-05-31 1965-05-11 Rolls Royce Axial flow compressor or fan and gas turbine engine provided therewith
US3191886A (en) * 1957-01-29 1965-06-29 Bristol Siddeley Engines Ltd Gas turbine jet propulsion engine units for aircraft and also jet propelled aircraft

Family Cites Families (3)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
BE463635A (en) * 1944-01-31
US2912188A (en) * 1955-09-15 1959-11-10 Bell Aircraft Corp Jet propelled aircraft with tiltable combustion chambers
DE1131949B (en) * 1958-09-05 1962-06-20 Gen Electric Runner for a two-circuit turbo jet engine

Patent Citations (4)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
US2509890A (en) * 1945-03-13 1950-05-30 Edward A Stalker Jet-propelled aircraft with boundary layer control
US3191886A (en) * 1957-01-29 1965-06-29 Bristol Siddeley Engines Ltd Gas turbine jet propulsion engine units for aircraft and also jet propelled aircraft
US3153906A (en) * 1960-12-15 1964-10-27 Bristol Siddeley Engines Ltd Aircraft propulsion power units
US3182898A (en) * 1962-05-31 1965-05-11 Rolls Royce Axial flow compressor or fan and gas turbine engine provided therewith

Cited By (2)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
US3545209A (en) * 1967-06-19 1970-12-08 Rolls Royce Gas turbine by-pass engine
US3867813A (en) * 1971-10-05 1975-02-25 Mtu Muenchen Gmbh Turbojet engine for vertical or short take-off and landing airplanes

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DE1264871B (en) 1968-03-28

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