GB2302371A - Gas turbine engine cooling air system - Google Patents

Gas turbine engine cooling air system Download PDF

Info

Publication number
GB2302371A
GB2302371A GB9512660A GB9512660A GB2302371A GB 2302371 A GB2302371 A GB 2302371A GB 9512660 A GB9512660 A GB 9512660A GB 9512660 A GB9512660 A GB 9512660A GB 2302371 A GB2302371 A GB 2302371A
Authority
GB
United Kingdom
Prior art keywords
gas turbine
turbine engine
cooling air
engine
casing
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.)
Withdrawn
Application number
GB9512660A
Other versions
GB9512660D0 (en
Inventor
Arnold Charles Newton
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.)
Rolls Royce PLC
Original Assignee
Rolls Royce PLC
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 Rolls Royce PLC filed Critical Rolls Royce PLC
Priority to GB9512660A priority Critical patent/GB2302371A/en
Publication of GB9512660D0 publication Critical patent/GB9512660D0/en
Publication of GB2302371A publication Critical patent/GB2302371A/en
Withdrawn legal-status Critical Current

Links

Classifications

    • B—PERFORMING OPERATIONS; TRANSPORTING
    • B64—AIRCRAFT; AVIATION; COSMONAUTICS
    • B64D—EQUIPMENT FOR FITTING IN OR TO AIRCRAFT; FLIGHT SUITS; PARACHUTES; ARRANGEMENT OR MOUNTING OF POWER PLANTS OR PROPULSION TRANSMISSIONS IN AIRCRAFT
    • B64D29/00—Power-plant nacelles, fairings, or cowlings
    • B64D29/02—Power-plant nacelles, fairings, or cowlings associated with wings
    • B—PERFORMING OPERATIONS; TRANSPORTING
    • B64—AIRCRAFT; AVIATION; COSMONAUTICS
    • B64C—AEROPLANES; HELICOPTERS
    • B64C21/00—Influencing air flow over aircraft surfaces by affecting boundary layer flow
    • B64C21/02—Influencing air flow over aircraft surfaces by affecting boundary layer flow by use of slot, ducts, porous areas or the like
    • B64C21/04—Influencing air flow over aircraft surfaces by affecting boundary layer flow by use of slot, ducts, porous areas or the like for blowing
    • B—PERFORMING OPERATIONS; TRANSPORTING
    • B64—AIRCRAFT; AVIATION; COSMONAUTICS
    • B64D—EQUIPMENT FOR FITTING IN OR TO AIRCRAFT; FLIGHT SUITS; PARACHUTES; ARRANGEMENT OR MOUNTING OF POWER PLANTS OR PROPULSION TRANSMISSIONS IN AIRCRAFT
    • B64D33/00—Arrangements in aircraft of power plant parts or auxiliaries not otherwise provided for
    • B64D33/08—Arrangements in aircraft of power plant parts or auxiliaries not otherwise provided for of power plant cooling systems
    • B—PERFORMING OPERATIONS; TRANSPORTING
    • B64—AIRCRAFT; AVIATION; COSMONAUTICS
    • B64C—AEROPLANES; HELICOPTERS
    • B64C2230/00—Boundary layer controls
    • B64C2230/04—Boundary layer controls by actively generating fluid flow
    • B—PERFORMING OPERATIONS; TRANSPORTING
    • B64—AIRCRAFT; AVIATION; COSMONAUTICS
    • B64D—EQUIPMENT FOR FITTING IN OR TO AIRCRAFT; FLIGHT SUITS; PARACHUTES; ARRANGEMENT OR MOUNTING OF POWER PLANTS OR PROPULSION TRANSMISSIONS IN AIRCRAFT
    • B64D33/00—Arrangements in aircraft of power plant parts or auxiliaries not otherwise provided for
    • B64D33/02—Arrangements in aircraft of power plant parts or auxiliaries not otherwise provided for of combustion air intakes
    • B64D2033/024—Arrangements in aircraft of power plant parts or auxiliaries not otherwise provided for of combustion air intakes comprising cooling means
    • F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F05—INDEXING SCHEMES RELATING TO ENGINES OR PUMPS IN VARIOUS SUBCLASSES OF CLASSES F01-F04
    • F05B—INDEXING SCHEME RELATING TO WIND, SPRING, WEIGHT, INERTIA OR LIKE MOTORS, TO MACHINES OR ENGINES FOR LIQUIDS COVERED BY SUBCLASSES F03B, F03D AND F03G
    • F05B2260/00—Function
    • F05B2260/20—Heat transfer, e.g. cooling
    • Y—GENERAL 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
    • Y02—TECHNOLOGIES OR APPLICATIONS FOR MITIGATION OR ADAPTATION AGAINST CLIMATE CHANGE
    • Y02T—CLIMATE CHANGE MITIGATION TECHNOLOGIES RELATED TO TRANSPORTATION
    • Y02T50/00—Aeronautics or air transport
    • Y02T50/10—Drag reduction
    • Y—GENERAL 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
    • Y02—TECHNOLOGIES OR APPLICATIONS FOR MITIGATION OR ADAPTATION AGAINST CLIMATE CHANGE
    • Y02T—CLIMATE CHANGE MITIGATION TECHNOLOGIES RELATED TO TRANSPORTATION
    • Y02T50/00—Aeronautics or air transport
    • Y02T50/60—Efficient propulsion technologies, e.g. for aircraft

Landscapes

  • Engineering & Computer Science (AREA)
  • Aviation & Aerospace Engineering (AREA)
  • Chemical & Material Sciences (AREA)
  • Combustion & Propulsion (AREA)
  • Mechanical Engineering (AREA)
  • Turbine Rotor Nozzle Sealing (AREA)

Abstract

In a gas turbine engine 10 cooling air which has been utilised to cool accessories 20 is ducted to an exit which directs the air over the engine casing structure so as to re-energise boundary layer flow and/or reduce base drag. The exit may comprise apertures (38, 40) formed in the fillets at the junction between an engine casing and the casing of a pylon supporting the engine (figure 2), or slot-like outlets (64) formed at the trailing end of a support pylon casing (figure 3).

Description

GAS TURBINE ENGINE WITH IMPROVED COOLING AIR SYSTEM The present invention relates to a gas turbine engine which may have a ducted fan, and does include an apparatus cooling airflow system.
Particularly but not restrictively, the invention relates to a gas turbine engine in which a cooling airflow is induced for the purpose of cooling apparatus such as oil tanks, oil coolers, gearboxes, fuel pumps and the like, which are used on gas turbine engines that power aircraft.
It is the practice, to dump any induced gas flow other than engine exhaust gas flow, from any convenient point on the engine carcass. By "dump" is meant, to eject overboard without consideration of potential uses over and above that for which the gas flow was induced.
The present invention seeks to provide a gas turbine engine including an improved cooling air exhaust system.
According to the present invention a gas turbine engine comprises casings defining closed spaces, apparatus within a said space, ducting extending from a cooling air inlet in one of said casings to and about said apparatus and thereafter to at least one used cooling air outlet in at least one of said casings, which air outlet is so formed as to cause used cooling air passing therefrom to re-energise local boundary layer flow and/or reduce base drag on an external surface of said at least one of said casings.
The invention will now be described, by way of example and with reference to the accompanying drawings in which: Fig 1 is a diagrammatic, axial cross sectional view of a gas turbine engine in accordance with the present invention.
Fig 2 is a view in the direction of arrow 2 in Fig 1.
Fig 3 is a pictorial view of the engine in Fig 1 but depicts a further embodiment of the present invention.
Referring to Fig 1. A gas turbine ducted fan engine 10 has a streamlined casing 12 surrounding a core gas generator 14. The casing 12 and the non-streamlined casing 16 of the core gas generator 14 are radially spaced and the space defined thereby is indicated by the numeral 18.
Apparatus which in the present example is an air cooled oil cooler 20, is supported in the space 19 by any suitable, known means An air inlet 22 is located in the casing 12, at a position upstream of the apparatus 20. By upstream is meant, with respect to the direction of flow of gases through the engine 10.
In the present example, the casing 12 is surrounded by and connected to a fan casing 24, the connection being made via a further casing 26 which encloses a pylon structure 28 in known manner. The pylon 28 is in turn connected between the engine 10 and a beam (not shown) in the wing 30 of the associated aircraft, again in known manner.
The casing 12 is faired into the pylon casing 26 by way of curved fillet pieces 32, 34, which again per se is a known feature and is best seen in Fig 2.
Ducting 36 extends from the air inlet 22, to the space 18 and surrounds the apparatus 20, before dividing, to connect with respective air outlets 38,40 formed one in each fillet 32,34 such that they will discharge used cooling air along the downstream portion of their respective fillet 32,34, ie in the same direction as the flow of fan air over the casing, 12, during flight of an associated aircraft. Thus, fan airflow will not be disturbed, and the boundary layer will be re-energised, having slowed through friction between itself and the casing 12.
The fan casing 24 is also faired into the casing 26, and if desired, the ducting 36 can be extended to connect with outlets 47,49 so as to achieve the same benefits (Fig 2).
If it is desired to cool the turbine casing 42 of the core gas generator 14, after cooling the apparatus 20 and prior to ejecting the air so used overboard, a plenum chamber 44 must be provided, eg by the casings 12 and 14 and bulk-heads 46,48. At least some of the used air will then circulate around the casing 42 prior to passing to atmosphere via outlets 38,40, thus reducing localised thermal effects.
A more efficient mode of cooling casing 42 would be to obviate air outlets 38,40 and substitute others (43,45) at positions diametrically opposite the fillets 32,34. By this means all of the cooling air, after passing around the apparatus 20, will flow around the casing 42 prior to leaving the newly positioned exits (not shown).
A further embodiment of the present invention and superimposed on Fig 1 provides apparatus 20 in the pylon casing 26. Ducting 50 extends into the space 52 in which the apparatus 20 resides, with fan air inlets formed in each side of the casing 26. Only one inlet, 54 is shown.
The ducting encloses the apparatus 20 and then extends further, to connect with air outlets on each side of the casing 26, near the downstream edge 56 thereof, which in this example is formed into a knife edge 58 which is best seen in Fig 2. Only one air outlet 60 is shown, but both are so formed as to ensure that air passing therefrom will flow along the side surfaces of the casing 26 and thus will re-energise the boundary layer in that area.
Alternatively, the ducting 50 could be extended to the outlets 38,40. The expert in the field, on reading this specification, will appreciate that either or both pieces of apparatus 20 could be incorporated in their respective locations, and air inlets 22,54, outlets 38,40 and/or 60 adopted, depending on design requirements.
Whilst the fan casing 26 of Figs 1 and 2 terminates in a knife edge 58, in Fig 3 to which reference is now made, the casing 26a terminates in a flat surface 62 which lies in a plane laterally of the engine axis. This arrangement causes turbulence in the fan air as it leaves the edges of the flat surface 62 which then generates a low pressure area at that surface. The term of art for this phenomenon is base drag. In order that this phenomenon be countered, slot type air outlets 64 are formed in the flat surface 62 such that their long dimensions span it. The outlets 64 are so formed as to cause used cooling air which passes therethrough, to flow along the flat surface and thus exert a static pressure thereon which at least reduces the base drag effect.
A further benefit which is derived from the embodiment of Fig 3, is that on losing its initial ejection velocity, the outlet air is entrained by the fan air leaving the edges of the flat surface 62 with the result that turbulence there is virtually eliminated.
No ducting or apparatus 20 is shown in Fig 3, the expert realising that the arrangements shown in Figs 1 and 2 as being capable of straightforward transfer to Fig 3, except that the outlet portions of the ducts must be re-aligned and divided in a manner depending on the number of slot outlets 64 incorporated. Alternatively, the ducts could terminate in a plenum chamber (not shown) which would provide a common supply to all of the slots.
Whilst the invention has been described with particular reference to a ducted fan gas turbine engine, which is supported from a pylon, it is equally applicable to a gas turbine engine per se, in which case the air inlets and outlets would be formed in the outer casing of the engine. Alternatively it may be convenient to form the outlets in any hollow flow straightening vanes at the downstream end of the engine.
The invention may also be applied to an engine which is buried in an aircraft wing or fuselage. In these cases, the wing "skin" or fuselage "skin" would provide the location for the air inlets and outlets.

Claims (12)

Claims:
1. A gas turbine engine comprising casings defining closed spaces, apparatus within a said space, ducting extending from a cooling air inlet in one of said casings to and about said apparatus and thereafter to at least one used cooling air outlet in at least one of said casings, which air outlet is so formed as to cause used cooling air passing therefrom to re-energise local boundary layer flow and/or reduce base drag on an external surface of said at least one of said casings.
2. A gas turbine engine as claimed in claim 1 comprising nested, spaced casings, apparatus within said space, ducting leading from a cooling air inlet in the outer one of said casings to and about said apparatus and thereafter to at least one used cooling air exit in said outer casing, which air exit is so formed as to cause used cooling air passing therefrom, to re-energise local boundary layer flow and/or reduce base drag.
3. A gas turbine engine as claimed in claim 2 wherein at least the inner one of said nested casings is of at least generally cylindrical cross-sectional shape.
4. A gas turbine engine as claimed in claim 3 wherein said engine is adapted for connecting to an aircraft via a pylon which is surrounded by a casing which connects via curved fillets with said engine outer casing, and said at least one used cooling air outlet is formed in at least one of said fillets.
5. A gas turbine engine as claimed in claim 4 wherein two used cooling air outlets are provided and reside, one in each curved fillet.
6. A gas turbine engine as claimed in claim 1 to 6 wherein said engine is adapted for connection to an aircraft via a pylon surrounded by a casing and the downstream end of said pylon casing defines a flat surface which lies in a plane laterally of the engine axis and said at least one used cooling air exit is defined by a slot in said flat surface, which slot is so arranged and shaped, as to cause used cooling air passing therefrom to flow over said flat surface and so fill the void which in operation of the engine is caused by ambient airflow breaking away from the edges of said flat surface, and thereby reduce base drag.
7. A gas turbine engine as claimed in claim 6 wherein a plurality of said slots are provided.
8. A gas turbine engine as claimed in any of claims 1 to 3 wherein said outer one of said nested casings comprises an aircraft wing, said air entry and exit being formed therein at leading and trailing wing portions respectively, both said portions defining a part of the upper wing skin.
9. A gas turbine engine as claimed in any previous claim wherein the engine includes a ducted fan.
10. A gas turbine engine substantially as described in this specification.
11. A gas turbine engine substantially as described in this specification with respect to Figs 1 and 2 of the drawings.
12. A gas turbine engine substantially as described in this specification with respect to Fig 3 of the drawings.
GB9512660A 1995-06-21 1995-06-21 Gas turbine engine cooling air system Withdrawn GB2302371A (en)

Priority Applications (1)

Application Number Priority Date Filing Date Title
GB9512660A GB2302371A (en) 1995-06-21 1995-06-21 Gas turbine engine cooling air system

Applications Claiming Priority (1)

Application Number Priority Date Filing Date Title
GB9512660A GB2302371A (en) 1995-06-21 1995-06-21 Gas turbine engine cooling air system

Publications (2)

Publication Number Publication Date
GB9512660D0 GB9512660D0 (en) 1995-08-23
GB2302371A true GB2302371A (en) 1997-01-15

Family

ID=10776453

Family Applications (1)

Application Number Title Priority Date Filing Date
GB9512660A Withdrawn GB2302371A (en) 1995-06-21 1995-06-21 Gas turbine engine cooling air system

Country Status (1)

Country Link
GB (1) GB2302371A (en)

Cited By (10)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
FR2891255A1 (en) * 2006-07-07 2007-03-30 Airbus France Sas Engine e.g. jet engine, assembly for aircraft, has outlet placed in rear with respect to rear engine mount, and heat exchanger system with exchanger arranged inside fairing that is entirely situated in rear with respect to engine mount
FR2891248A1 (en) * 2005-09-28 2007-03-30 Airbus France Sas Engine e.g. jet engine, assembly for aircraft, has heat exchange system with outlet situated between shell and engine and rearwards with respect to rear engine attachment interposed between engine and engine mounting structure
GB2438696A (en) * 2006-04-20 2007-12-05 Rolls Royce Plc Exhaust outlet for a turbine engine heat exchanger
DE102009010647A1 (en) * 2009-02-26 2010-09-02 Rolls-Royce Deutschland Ltd & Co Kg Running column adjustment system of an aircraft gas turbine
FR2961173A1 (en) * 2010-06-09 2011-12-16 Airbus Operations Sas NACELLE INCORPORATING AN AIR INLET AT A CAP
FR2977567A1 (en) * 2011-07-07 2013-01-11 Airbus Operations Sas METHOD FOR COOLING A THERMAL PROTECTION FLOOR OF AERODYNAMIC REAR FITTING OF A FITTING MAT OF A PROPELLANT AIRCRAFT ASSEMBLY
FR3001199A1 (en) * 2013-01-23 2014-07-25 Snecma MOTOR COVER INCORPORATING AN EQUIPMENT VENTILATION CIRCUIT
FR3013330A1 (en) * 2013-11-20 2015-05-22 Snecma AIRCRAFT COMPRISING AN OIL TANK DEPORTE
DE102017117291A1 (en) * 2017-07-31 2019-01-31 Rolls-Royce Deutschland Ltd & Co Kg Air guiding device in an aircraft engine
CN110159358A (en) * 2018-02-14 2019-08-23 中国航发商用航空发动机有限责任公司 Casing between grade

Cited By (21)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
FR2891248A1 (en) * 2005-09-28 2007-03-30 Airbus France Sas Engine e.g. jet engine, assembly for aircraft, has heat exchange system with outlet situated between shell and engine and rearwards with respect to rear engine attachment interposed between engine and engine mounting structure
WO2007036523A1 (en) * 2005-09-28 2007-04-05 Airbus France Engine assembly for an aircraft comprising an engine as well as an engine mounting structure for such an engine
WO2007036522A1 (en) * 2005-09-28 2007-04-05 Airbus France Engine assembly for an aircraft comprising an engine as well as an engine mounting structure for such an engine
US7971826B2 (en) 2005-09-28 2011-07-05 Airbus Operations Sas Engine assembly for an aircraft comprising an engine as well as an engine mounting structure for such an engine
US8061649B2 (en) 2005-09-28 2011-11-22 Airbus Operations Sas Engine assembly for an aircraft comprising an engine as well as an engine mounting structure for such an engine
GB2438696A (en) * 2006-04-20 2007-12-05 Rolls Royce Plc Exhaust outlet for a turbine engine heat exchanger
GB2438696B (en) * 2006-04-20 2009-10-07 Rolls Royce Plc A gas turbine engine
US7886520B2 (en) 2006-04-20 2011-02-15 Rolls-Royce Plc Gas turbine engine
FR2891255A1 (en) * 2006-07-07 2007-03-30 Airbus France Sas Engine e.g. jet engine, assembly for aircraft, has outlet placed in rear with respect to rear engine mount, and heat exchanger system with exchanger arranged inside fairing that is entirely situated in rear with respect to engine mount
US8834108B2 (en) 2009-02-26 2014-09-16 Rolls-Royce Deutschland Ltd & Co Kg Running-gap control system of an aircraft gas turbine
DE102009010647A1 (en) * 2009-02-26 2010-09-02 Rolls-Royce Deutschland Ltd & Co Kg Running column adjustment system of an aircraft gas turbine
FR2961173A1 (en) * 2010-06-09 2011-12-16 Airbus Operations Sas NACELLE INCORPORATING AN AIR INLET AT A CAP
US9067686B2 (en) 2010-06-09 2015-06-30 Airbus Operations (Sas) Nacelle comprising an air entrance in a cap
EP2543864A3 (en) * 2011-07-07 2014-03-26 Airbus Operations (Société par actions simplifiée) Aircraft propulsion assembly with a heat shield for thermal protection of a rear aerodynamic fairing of a pylon and a cooling method for the heat shield
FR2977567A1 (en) * 2011-07-07 2013-01-11 Airbus Operations Sas METHOD FOR COOLING A THERMAL PROTECTION FLOOR OF AERODYNAMIC REAR FITTING OF A FITTING MAT OF A PROPELLANT AIRCRAFT ASSEMBLY
US9435224B2 (en) 2011-07-07 2016-09-06 Airbus Operations S.A.S. Method for cooling a thermal protection floor of an aft aerodynamic fairing of a structure for mounting an aircraft propulsion system
FR3001199A1 (en) * 2013-01-23 2014-07-25 Snecma MOTOR COVER INCORPORATING AN EQUIPMENT VENTILATION CIRCUIT
US10077113B2 (en) 2013-01-23 2018-09-18 Safran Aircraft Engines Engine cowl incorporating an equipment ventilation circuit
FR3013330A1 (en) * 2013-11-20 2015-05-22 Snecma AIRCRAFT COMPRISING AN OIL TANK DEPORTE
DE102017117291A1 (en) * 2017-07-31 2019-01-31 Rolls-Royce Deutschland Ltd & Co Kg Air guiding device in an aircraft engine
CN110159358A (en) * 2018-02-14 2019-08-23 中国航发商用航空发动机有限责任公司 Casing between grade

Also Published As

Publication number Publication date
GB9512660D0 (en) 1995-08-23

Similar Documents

Publication Publication Date Title
US4466587A (en) Nacelle installation
EP0626039B1 (en) Exhaust eductor cooling system
US5947412A (en) Jet engine noise suppressor assembly
US7716913B2 (en) Engine
US5014933A (en) Translating lip aircraft cowling structure adapted for noise reduction
US20220242583A1 (en) Aircraft with an offset nacelle aligned with the wake of the wing
US4930725A (en) Pusher propeller installation for turboprop engines
EP2204567B1 (en) Apparatus and method for controlling the boundary layer in a gas turbine engine
US4892269A (en) Spinner ducted exhaust for pusher turboprop engines
EP3171009B1 (en) Compression cowl for jet engine exhaust
US5653110A (en) Film cooling of jet engine components
CA2926862A1 (en) Turbofan bypass air cooled oil cooler fairings
JPH07500169A (en) Turbojet engine with fan or prop fan
GB2302371A (en) Gas turbine engine cooling air system
CA1263242A (en) Gas turbine outlet arrangement
US10556702B2 (en) Aircraft having an airflow duct
US8403629B2 (en) Stator for a jet engine, a jet engine comprising such a stator, and an aircraft comprising the jet engine
US20090217643A1 (en) Gas discharge device for a vehicle engine
GB2211247A (en) Turboprop exhaust nozzle
GB2138507A (en) Mounting end exhausting in turbo-propellor aircraft engines
US3900177A (en) Jet propulsion powerplant
US10974813B2 (en) Engine nacelle for an aircraft
US11772779B2 (en) Propulsion unit with improved boundary layer ingestion
EP0440640A1 (en) Spinner ducted exhaust for pusher turboprop engines
Kong Propulsion system Integration of turboprop aircraft for basic trainer

Legal Events

Date Code Title Description
WAP Application withdrawn, taken to be withdrawn or refused ** after publication under section 16(1)