• 专利附录
  • 专利说明
  • 权利要求
  • 类似技术
  • 同族专利
  • 引用文献
  • 法律状态
  • 优先权
    • 专利摘要:
    An aircraft turbomachine (110) comprising a surface heat exchanger (50) comprising an oil circuit and at least one heat exchange surface with the oil and intended to be swept by a flow of air, a duct (52) for circulating said air flow in which said exchanger is mounted and comprising an air inlet (54) and an air outlet (56), means (44) for discharging air from a vein compressor, characterized in that said discharge means comprises an air outlet connected to said duct so that discharge air can be delivered to said exchanger.
    公开号:FR3044636A1
    申请号:FR1561981
    申请日:2015-12-08
    公开日:2017-06-09
    发明作者:Nicolas Claude Parmentier；Nicolas Maurice Herve Aussedat
    申请人:SNECMA SAS；
    IPC主号:
    专利说明:

    Aircraft turbomachine equipped with an air-oil surface heat exchanger
    TECHNICAL AREA
    The present invention relates to an aircraft turbine engine, of the type comprising an air-oil surface heat exchanger.
    STATE OF THE ART The state of the art includes in particular document FR 2 788 308.
    An aircraft turbomachine conventionally comprises a gas engine or generator comprising, from upstream to downstream, in the direction of flow of the gases in the turbomachine, at least one compressor, a combustion chamber, and at least one turbine. The engine is supplied with air by an air intake sleeve and a nozzle makes it possible to evacuate the combustion gases leaving the engine turbine.
    In the case of a twin-engine turbomachine, its engine comprises a compressor and a low pressure turbine or LP whose rotors are connected together by a LP shaft and form with this shaft a BP body, and a compressor and a high pressure turbine or HP whose rotors are connected together by an HP shaft and form with this tree an HP body. The engine thus comprises an air flow (so-called primary) flow intended to feed the combustion chamber, and a flow stream of combustion gas leaving the chamber.
    In the case of a blower turbine engine, a blower is mounted in the air inlet duct (and is therefore faired) and is driven by the LP shaft, either directly or indirectly via a control box. Gear reducer gears.
    In the case of a turboprop, a non-faired external propeller is driven by the LP shaft or a free turbine, through a reduction gearbox. In general, the entry of air into a turboprop engine is through the front and the exit of the combustion gases is from the rear. However, an inverted type turboprop engine is known, in which the air inlet is via the rear and the outlet of the combustion gases is from the front.
    It is known to equip a turbomachine with an air-oil surface heat exchanger, also called ACOC (acronym for English Air Cooled OR Cooler). This exchanger is generally mounted in a flow duct of an air flow having an air inlet and an air outlet. The exchanger comprises at least one heat exchange surface with the oil and intended to be swept by the air flow.
    In some cases of use, the flow of air does not circulate naturally or sufficiently in the conduit. Assistance involving additional sampling on the turbomachine may therefore be necessary. In the case of a turboprop, a sampling in the compressor vein can be envisaged at low speed to force the flow of a sufficient air flow on the exchanger, for example by means of a fan or a fan. a jet pump. In the case of a turbofan engine, sampling via a bailer in the flow vein of the secondary stream is possible to divert part of the secondary flow to the exchanger.
    Air discharge means, also called means of operability, which are generally valves (VBV which is the acronym for Variable Bleed Valve or HBV which is the acronym for Handling Bleed Valve) have the function of evacuating air to the outside of a vein (primary or combustion gas ejection), which is a loss of pneumatic energy.
    These discharge means are necessary to ensure the proper functioning of the engine especially at idle or partial regimes (between idle and full throttle for each phase of flight). This loss of pneumatic energy is often expensive in all cases: - in the case of a HBV valve of an existing first engine: the discharge is made in the HP compressor. The air has been heavily compressed which represents a high power lost despite an average discharge rate; - in the case of a HBV valve of a second existing engine or a VBV valve of a third existing engine: the discharge is made just behind a compressor BP or intermediate, in which the air is less compressed than in the previous case. On the other hand, the unloaded flow is greater than on a HBV valve taken in an HP compressor. The compression power of this discharged air thus remains high.
    On the second engine mentioned above, the power consumed by the compression of the air passing through a HBV valve can reach 2.5% of the maximum power of the engine in certain cases of emergency flight (for example in the case of stopping one of the two engines). In general, the power consumption by the discharge means can be up to 30% of the flow rate through the turbomachine.
    In general, a turbomachine is optimized for a point of operation in full gas. This implies that in partial or idle mode the operation of the turbomachine is not optimized. In particular, it is necessary to restore operating margins of the turbomachine using the air discharge system at one of the compressors as previously described. This air bleed has the effect of causing an increase in the temperatures of the gases at the output of the turbomachine at idle. Idle gas exit temperatures are potentially higher than full gas outlet temperatures. This poses two problems: it is necessary to have a free turbine, an outlet casing, a gas ejection nozzle made of materials capable of withstanding these temperatures; these materials have the disadvantages of being expensive and heavy; to overcome this problem, in a continuing effort to save weight and cost of construction, cooling equipment has been put in place to limit the temperature of these gases; this equipment is penalizing in terms of mass, even if its mass ratio compared to thermo-resistant materials is positive; moreover, this equipment is penalizing in terms of congestion; - On the other hand, during its use, the turbomachine spends more time idling than full gas; this has the effect of degrading the engine by significant wear of parts (resulting in significant maintenance costs).
    The present invention provides an improvement to existing technologies.
    SUMMARY OF THE INVENTION The invention proposes an aircraft turbomachine, comprising: a surface heat exchanger comprising an oil circuit and at least one heat exchange surface with the oil and intended to be swept by a flow of air, - a circulation duct of said air flow in which said exchanger is mounted and comprising an air inlet and an air outlet, - air discharge means of a compressor stream, characterized in that, said discharge means comprising a discharge air outlet connected to said duct, the discharge air is supplied to said exchanger.
    The present invention thus proposes a pooling of the operability and thermal management functions of the air-oil surface exchanger. It proposes the use of a discharge as a source of pneumatic energy to assist the cooling of the exchanger. The invention makes it possible to introduce the concept of pneumatic energy recovery, which has never been seen on a standard aeronautical engine. This allows to bring an additional energy supply and not used before (which would be lost anyway), which improves the energy balance of the engine. In particular, the discharge and cooling functions of the exchanger, made with pneumatic energy, are pooled. This reduces the overall cost by re-optimization of all the systems and in particular the exchanger.
    The turbomachine according to the invention may comprise one or more of the following characteristics, taken separately from one another or in combination with each other: said air outlet opens upstream of said exchanger with reference to the flow of the air in said duct, - said inlet and outlet of the duct open on an outer surface of a nacelle of the turbomachine, - said inlet and outlet duct open on an outer surface of a casing delimiting the inner periphery of a ring vein of the turbomachine, the inlet of said duct is equipped with a scoop and / or a tilting door, said discharge means comprise a pipe which extends between a compressor stream and said duct, the turbine engine being a turboprop engine, - the turbomachine being a turbofan engine, - said surface heat exchanger is the only air-oil surface heat exchanger of the turbomachine.
    DESCRIPTION OF THE FIGURES The invention will be better understood and other details, characteristics and advantages of the invention will emerge more clearly on reading the following description given by way of nonlimiting example and with reference to the appended drawings in which: FIG. 1 is a schematic perspective view of a turbomachine which is here an inverted type turboprop, FIG. 2 is a very diagrammatic view in axial section of a turboprop of the type of FIG. 1, FIG. 3 is a very diagrammatic view in axial section of a turboprop of the type of FIG. 1, and illustrates the prior art to the present invention; FIG. 4 is a very schematic view in axial section of a turboprop of the following type; of FIG. 1, and illustrates the invention; FIG. 5 is a very schematic view in axial section of a turbine engine of the turbofan type, and illustrates the prior art technique; FIG. 6 is a schematic view in axial section of a turbomachine, and illustrates the invention.
    DETAILED DESCRIPTION
    Referring firstly to Figure 1 which shows an aircraft turbine engine and more particularly a first turboprop 10 which is here of the inverted type, although the invention is not limited to this particular application.
    The first turboprop 10 may comprise a motor 12 of longitudinal axis A and comprising a compressor 14, 16, an annular combustion chamber 18, and at least one turbine 20. The axis A corresponds to the axis of rotation of the rotors of the motor 12 and in particular compressor 14, 16 and turbine 20. The compressor 14, 16 may comprise a high pressure compressor 16 and a low pressure compressor 14. The at least one turbine 20 may be a low or high pressure turbine .
    The rotors of the high pressure compressor 16 and the high pressure turbine 20 are connected to each other by a high pressure shaft or HP centered on the axis A, and form a high pressure body or HP.
    As indicated in the foregoing, the illustrated example is non-limiting and the turboprop can be indifferently of the single-barrel type and linked turbine type dual body and linked turbine type double or triple body and free turbine, etc.. The nature of the internal architecture of the turbomachine is of little importance vis-à-vis the system considered in the invention.
    The first turboprop 10 further comprises, at the front of the engine 12, a gearbox 24, an input shaft is driven by a free turbine 25 and an output shaft drives an external propeller 26 of the turboprop. Gearbox 24 is known as PGB, which stands for Power Gear Box.
    The low pressure compressor 14 is supplied with air by an air inlet casing 28 which is itself connected to an air intake sleeve 30. The turbine 20 is connected to a casing 32 for exhausting the combustion gases, which is itself connected to an exhaust nozzle 34.
    The first turboprop 10 as shown in FIG. 1 is of the inverted type: the low-pressure compressor 14 is situated at the rear of the engine and the free turbine 25 is located at the front of the engine, that is to say the side of the gearbox 24 and the propeller 26. This is particularly advantageous in that the free turbine 25, shown in Figure 2, is connected directly to the gearbox 24, without the need for a shaft BP crossing the HP body.
    The nozzle 34 may be arranged on one or more sides of the engine (for example at 3 o'clock or 9 o'clock, by analogy with the dial of a clock). It comprises a gas inlet opening into the housing 32 and a gas outlet port 36 opening on one side of the turboprop, in the vicinity of its front end. The nozzle 34 may have in section a parallelepipedal shape elongated in a direction substantially perpendicular to the axis A, here substantially vertical. The nozzle defines a conduit for passing a flow of gas, called a second gas flow or hot flow.
    The air inlet sleeve 30 is for example preferably arranged under the engine (at 6 o'clock). It has an elongated shape, its axis of elongation being substantially parallel to the axis A. It extends over substantially the entire longitudinal dimension of the engine and comprises an air inlet orifice located at the front of the turboprop and an air outlet opening opening into the casing 28 of air inlet. The inlet duct 30 has for example in section a parallelepipedal shape elongated in a direction substantially perpendicular to the axis A, here substantially horizontal. The air intake sleeve defines a conduit for the passage of a flow of gas, called the first gas flow or cold flow.
    The primary flow supplying the engine comprises the cold flow entering and flowing into the air inlet sleeve 30, and the hot flow flowing into and out of the nozzle 34.
    The first turboprop engine 10 comprises means for discharging the primary flow air, which are for example discharge valves 40, as shown in FIG. 2. Each valve comprises a pivoting door, which is movable between a first position in which it closes a discharge port 41 and a second position in which it leaves free this orifice. When the gate of the valve 40 is in the second position, air is discharged to the outside of the engine through the discharge port 41. These discharge means are used to discharge only air and not air. combustion gas. They are therefore mounted upstream of the combustion chamber 18 and most often close to a compressor, such as the compressor BP 14 or HP 16. The discharged air is discharged by means of a duct 44.
    Figure 3 illustrates the first prior art turboprop 10, also of the inverted type, which is equipped with a surface-type air / oil heat exchanger 50 or ACOC. This exchanger 50 is mounted in a conduit 52 which extends substantially longitudinally vis-à-vis the axis of the engine. The duct 52 comprises upstream an inlet 54 for supplying air to the duct 52, and downstream an outlet 56 for discharging air from the duct. In the example shown, the inlets 54 and 56 are located on the outer surface of a nacelle 58 of the first turboprop 10.
    In the prior art illustrated in FIG. 3, a jet pump 60 is mounted in the duct 52, downstream of the heat exchanger 50, and projects a flow of air intended to force the flow of air in the duct. 52 from its entry to its exit and therefore through the exchanger 50. The jet pump 60 is generally powered by specific or dedicated air sampling means on the engine.
    A disadvantage of this technical solution is that the sample to feed the jet pump 60 is made in addition to the discharge, which is penalizing. This results in particular energy losses.
    FIG. 4 illustrates a second turboprop 110 according to the invention, also of the inverted type, which is equipped with an air / oil heat exchanger 50 of the surface type or ACOC. This exchanger 50 is mounted in a conduit 52 which extends substantially longitudinally vis-à-vis the axis of the engine. The duct 52 comprises upstream an inlet 54 for supplying air to the duct 52, and downstream an outlet 56 for discharging air from the duct. In the example shown, the inlets 54 and 56 are located on the outer surface of the nacelle 58 of the second turboprop 110. The exchanger 50 is supplied by the air flow taken to the ambient in normal use, for example by means of a bailer provided at the inlet of the conduit 52. The conduit 44 of the discharge means has its end opposite to the discharge orifice, which is connected to the conduit 52 and opens upstream of the exchanger 50.
    At this outlet, a movable flap 62 can be mounted. In normal use, the flap 62 is closed and the discharge port also. At idle, the flap can be opened so that discharge air sweeps the exchanger 50. The discharge is for example carried out at the last stages of the compressor. The opening of the flap 62 makes it possible to favor the discharge air for the air supply of the exchanger.
    With a substantial cold flow supplying the conduit 52 through its inlet, even at idle, the heat exchange remains effective and therefore the size of the exchanger 50 is reduced. The jet pump and the associated sample can be deleted. The example shown is a turboprop with an inverted body but the invention can likewise be applied with the same advantages to a conventional turboprop engine.
    FIG. 5 illustrates a turbomachine 210 of the prior art, which is equipped with an air / oil heat exchanger 50 of the surface type or ACOC. This exchanger 50 is mounted in a conduit 52 which extends substantially longitudinally vis-à-vis the axis of the engine. The duct 52 comprises upstream an inlet 54 for supplying air to the duct 52, and downstream an outlet 56 for discharging air from the duct. In the example shown, the inlet 54 is located on a casing delimiting the outer periphery of the flow duct of the secondary flow, and the outlet 56 is located on an outer surface of the nacelle 58 of the turbomachine 210.
    In the prior art illustrated in FIG. 5, a jet pump 60 is mounted in the duct 52, downstream of the exchanger 50, and projects a flow of air intended to force the flow of air in the duct. 52 from its entry to its exit and therefore through the exchanger 50. The jet pump 60 is generally powered by specific or dedicated air sampling means on the engine.
    FIG. 6 illustrates a turboprop engine 310 according to the invention, which is equipped with an air / oil heat exchanger 50 of surface type or ACOC type. This exchanger 50 is mounted in a conduit 52 which extends substantially longitudinally vis-à-vis the axis of the engine. The duct 52 comprises upstream an inlet 54 for supplying air to the duct 52, and downstream an outlet 56 for discharging air from the duct. In the example shown, the inputs 54 and 56 are located on a housing defining the inner periphery of the flow of the secondary flow stream. The exchanger 50 is supplied with a flow of air taken from the secondary flow in normal use, for example by means of a scoop provided at the inlet of the duct 52. The pipe 44 of the discharge means has its end opposite to the discharge orifice, which is connected to the conduit 52 and opens upstream of the exchanger 50.
    A movable flap 62 may be mounted at the inlet of the duct 52. In normal use, the flap 62 is open and the discharge port closed. At idle, the discharge port may be opened to blow discharge air through the exchanger 50. The flap 62 is closed to favor the use of the discharge air to that of the secondary flow.
    With a large cold flow, even at idle, the heat exchange remains effective and therefore the size of the exchanger is reduced. In addition, the device can be placed on the entire periphery of the housing, further reducing the length and height of the exchanger. The jet pump and the sampling in the associated compressor can therefore be deleted.
    In the end, the exchanger 50 may be smaller because dimensioned in cases of partial regime where the discharge ports are closed, that is to say for intermediate regimes between the partial regime and other regimes. Indeed, the heat exchanger 50 will not be dimensioned for idling as was the case in the state of the art: at idle, the flow of the discharge ports will be very important and enough to supply the exchanger. The ecoped flow rate will be lower than in the existing thus the penalty on the lowered cycle: the output flow of the exchanger is reinjected directly into the secondary vein which generates less loss than for the system of the state of the art . In addition, there will be no sampling for the jet jet at idle, which avoids overheating and excess consumption associated in this phase of engine use.
    权利要求:
    Claims (9)
    [1" id="c-fr-0001]
    An aircraft turbomachine (110,310) comprising: a surface heat exchanger (50) comprising an oil circuit and at least one heat exchange surface with the oil and intended to be swept by a flow of oil; air, - a duct (52) for circulating said air flow in which said exchanger is mounted and comprising an air inlet (54) and an air outlet (56), - means (40, 41, 44 ) for discharging air from a compressor stream, characterized in that said discharge means comprises a discharge air outlet connected to said conduit, the discharge air being supplied to said exchanger.
    [2" id="c-fr-0002]
    2. A turbomachine (110, 310) according to the preceding claim, wherein said air outlet (44) opens upstream of said exchanger (50) with reference to the flow of air in said conduit (52).
    [3" id="c-fr-0003]
    3. Turbomachine (110, 310) according to claim 1 or 2, wherein said inlet and outlet (54, 56) of the conduit (52) open on an outer surface of a nacelle (58) of the turbomachine.
    [4" id="c-fr-0004]
    4. A turbomachine (110, 310) according to claim 1 or 2, wherein said inlet and outlet (54, 56) of the conduit (52) open on an outer surface of a housing defining the inner periphery of an annular vein of the turbomachine.
    [5" id="c-fr-0005]
    5. Turbomachine (110, 310) according to one of the preceding claims, wherein the inlet (54) of said duct (52) is equipped with a bailer and / or a door (62) tilting.
    [6" id="c-fr-0006]
    6. Turbomachine (10, 10 ') according to one of the preceding claims, wherein said discharge means comprise a pipe (44) extending between a compressor stream and said conduit (52).
    [7" id="c-fr-0007]
    7. Turbomachine (110) according to one of the preceding claims, the turbomachine being a turboprop.
    [8" id="c-fr-0008]
    8. Turbomachine (310) according to one of claims 1 to 6, the turbomachine being a turbofan engine.
    [9" id="c-fr-0009]
    9. Turbomachine (110, 310) according to one of the preceding claims, wherein said surface heat exchanger (50) is the only air-oil surface heat exchanger of the turbomachine.
    类似技术:
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    同族专利:
    公开号 | 公开日
    FR3044636B1|2019-08-30|
    引用文献:
    公开号 | 申请日 | 公开日 | 申请人 | 专利标题
    US5438823A|1990-12-21|1995-08-08|Rolls-Royce, Plc|Heat exchange apparatus for gas turbine fluids|
    US5269135A|1991-10-28|1993-12-14|General Electric Company|Gas turbine engine fan cooled heat exchanger|
    FR2788308A1|1999-01-07|2000-07-13|Snecma|COOLING DEVICE FOR A TURBOMACHINE SPEED REDUCER|
    EP2348210A1|2010-01-26|2011-07-27|Airbus Operations |Aircraft propulsion unit with a cooler installed at engine nacelle|
    US20130098067A1|2011-10-21|2013-04-25|Gabriel L. Suciu|Constant speed transmission for gas turbine engine|
    US20150247462A1|2012-09-28|2015-09-03|United Technologies Corporation|Gas turbine engine thermal management system for heat exchanger using bypass flow|EP3561279A1|2018-04-24|2019-10-30|Rolls-Royce plc|Gas turbine engine|
    FR3093765A1|2019-03-12|2020-09-18|Safran Aircraft Engines|OPTIMIZED TURBOMACHINE AIR-OIL HEAT EXCHANGER SYSTEM|
    US11092074B2|2017-12-13|2021-08-17|Rolls-Royce Plc|Bleed ejector|
    法律状态:
    2016-12-05| PLFP| Fee payment|Year of fee payment: 2 |
    2017-06-09| PLSC| Publication of the preliminary search report|Effective date: 20170609 |
    2017-11-21| PLFP| Fee payment|Year of fee payment: 3 |
    2018-09-14| CD| Change of name or company name|Owner name: SAFRAN AIRCRAFT ENGINES, FR Effective date: 20180809 |
    2018-11-27| PLFP| Fee payment|Year of fee payment: 4 |
    2019-11-20| PLFP| Fee payment|Year of fee payment: 5 |
    2020-11-20| PLFP| Fee payment|Year of fee payment: 6 |
    2021-11-18| PLFP| Fee payment|Year of fee payment: 7 |
    优先权:
    申请号 | 申请日 | 专利标题
    FR1561981|2015-12-08|
    FR1561981A|FR3044636B1|2015-12-08|2015-12-08|AIRCRAFT TURBOMACHINE EQUIPPED WITH AIR-OIL SURFACE HEAT EXCHANGER|FR1561981A| FR3044636B1|2015-12-08|2015-12-08|AIRCRAFT TURBOMACHINE EQUIPPED WITH AIR-OIL SURFACE HEAT EXCHANGER|
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