Aircraft equipped with a heating system for a turboshaft engine plenum

Abstract

The present invention relates to an aircraft (1) provided with a turbine engine (10) comprising a gas generator (15), the gas generator (15) comprising a compression assembly (20) supplying air to a combustion chamber (24) and a turbine assembly (25) supplied with gas by the combustion chamber (24), the aircraft (1) having a plenum (40) supplying air to the compression assembly (20). The aircraft (1) comprises a heating system (50), the heating system (50) comprising a heat exchanger (60) arranged in the plenum (40), the heating system (50) comprising a supply fluid connection (70) and a discharge fluid connection (80) connected to the heat exchanger (60), the supply fluid connection (70) conveying hot air from the turbine engine (10) into the heat exchanger (60). Abstract figure: figure 2

Description

Title of the invention: Aircraft equipped with a heating system for a turboshaft engine plenum

[0001] The present invention relates to an aircraft provided with a heating system for a turbine engine plenum. Such an aircraft may be a rotorcraft. The technical field of the invention therefore relates to the field of engine air supply systems.

[0002] In particular, an engine may be a turboshaft engine comprising a gas generator and at least one turbine. The gas generator is provided with a compression assembly supplying compressed air to a combustion chamber. In addition, the gas generator is provided with a turbine assembly set in motion by the hot gases leaving the combustion chamber. The turbine assembly is rotationally integral with the compression assembly.

[0003] To supply the compression assembly of the gas generator with air from outside, the aircraft comprises an air supply system. Depending on the architecture of the turbine engine and the aircraft, the air supply system may comprise a radial air inlet surface. A radial air inlet air supply system is also static, with air entering the radial air inlet surface being mainly drawn in by the turbine engine. Conversely, a dynamic air supply system comprises an air inlet ingesting air under the effect of the aircraft moving forward.

[0004] A radial air inlet air supply system comprises one or more inlet sections. Each inlet section is provided with a radial air inlet surface, generally rectangular, arranged radially relative to the turbine engine. This inlet section is connected by an annular pipe to the gas generator of the turbine engine. Such a system may be conventionally referred to as a "plenum" by those skilled in the art.

[0005] Thus, a plenum comprises an annular duct arranged around a central axis along which a turbine engine extends. The annular duct thus delimits an annular cavity in fluid communication with one or more inlet sections and the turbine engine.

[0006] Therefore, air outside the aircraft enters the plenum through a radial air inlet section and is then directed radially through the annular duct toward the turbine engine.

[0007] Optionally, a grid is placed at the interface between the plenum and the turbine engine to prevent the turbine engine from ingesting undue particles.

[0008] Document FR 3007798 A thus describes a plenum.

[0009] When the aircraft is flying in icing conditions, snow or ice may accumulate at the bottom of the plenum. The bottom of the plenum may be the lowest part of the plenum for predetermined permitted aircraft attitude angles. The plenum may include drains to evacuate the water contained in the plenum. The rolling movements of the aircraft may also allow snow or ice to be evacuated from the plenum. In addition, the engine is designed to operate normally following the ingestion of a certain quantity of snow or ice. It should be noted that the phenomenon encountered is in fact different on a dynamic air intake, in particular the problem of accretion of snow or ice on a bottom wall of the plenum since the air can enter a dynamic air intake at a high speed.

[0010] However, the ingestion of too much snow or ice is likely to damage the blades of the compression assembly of the turbine engine, or even to extinguish the combustion chamber in extreme cases. Therefore, the plenum is defined so that the turbine engine ingests a quantity of snow or ice below a threshold.

[0011] Some devices for combating the formation of ice and snow include electric heating mats. Such systems can be relatively complicated and require significant electrical energy.

[0012] Patent FR2924471 B1 discloses a filtration system potentially equipped with a means of heating a grid.

[0013] Patent EP2129579 B1 describes a dynamic air intake vent comprising a toroidal air circulation channel for heating the leading edge of this intake vent. This patent aims to prevent the formation of ice or snow and is therefore far removed from the problem of limiting ice or snow on the bottom of a plenum. In other words, patent EP2129579 B1 relates to an anti-icing system, not a de-icing system.

[0014] The present invention therefore aims to propose an aircraft equipped with an innovative system aimed at minimizing the quantity of snow and / or ice likely to be ingested by a turbine engine.

[0015] The present invention relates to an aircraft equipped with a turbine engine comprising a gas generator, the gas generator comprising a compression assembly supplying compressed air to a combustion chamber, the gas generator comprising a turbine assembly supplied with gas by the combustion chamber, the aircraft having a radial inlet air supply system equipped with a plenum supplying air to the compression assembly, the plenum comprising a duct, for example substantially annular, equipped with an external opening open to an external environment located outside the aircraft, or even an internal opening in communication fluidic with the turbine engine. The external opening may be a static radial opening, namely one that extends around a central axis along which the turbine engine is arranged. The central axis may be an axis of symmetry of at least a portion of the plenum and / or the turbine engine.

[0016] This aircraft comprises a heating system, the heating system comprising a heat exchanger arranged in a volume delimited by the plenum, the heating system comprising a supply fluid connection and a discharge fluid connection connected to the heat exchanger, the supply fluid connection conveying hot air from the turbine engine into the heat exchanger.

[0017] Flight in icing conditions can be problematic. Some aircraft may have a limited flight envelope in such conditions due to the risks associated with ice or snow being ingested into a turbine engine.

[0018] To solve this problem, the heat exchanger of the invention is supplied with hot air by the turbine engine under such conditions. This heat exchanger is thus heated by hot air and makes it possible to reach a skin temperature above zero degrees Celsius, or even 10 degrees Celsius. Therefore, the heat exchanger can heat the plenum by convection and radiation, and can thus limit the quantity of ice or snow in the plenum, in particular by melting it. This solution goes against prejudices since the plenum has the function of capturing fresh air from outside. Arranging a heating system in the plenum seems contrary to the ingestion of fresh air. Nevertheless, the use of the heating system in icing conditions can have an acceptable impact on the operation of the engine.The heating system of the invention therefore proves to be generally interesting in having a limited impact on the temperature of the air ingested by the turbine engine while reducing the quantity of ice or snow likely to be ingested into the turbine engine.

[0019] Furthermore, the heating system is simple and can be arranged on an existing aircraft.

[0020] The aircraft may further comprise one or more of the following features.

[0021] According to one possibility, the turbine engine comprising a gas stream extending from the plenum passing successively through the compression assembly then the combustion chamber and the turbine assembly, the fluid supply connection may be in fluid connection with the gas stream downstream of a compression stage of the compression assembly.

[0022] The fluid supply connection has a socket referred to by a person skilled in the art as a “P3 socket” to be supplied with hot air coming from the gas stream. The air taken is air at high temperature, for example of the order of 200 to 400 degrees Celsius, and under pressure, for example of the order of 6 to 9 bars, obtained by compression air ingested by the turbine engine.

[0023] Thus, the air taken in has a higher temperature than the ingested air and makes it possible to heat the heat exchanger, and consequently the plenum. The quantity of air taken in to supply the heat exchanger can be relatively low, for example of the order of 5 to 10 grams per second, and has no significant impact on the operation of the turbine engine.

[0024] The heating system can thus be relatively simple, unlike an electric heater requiring significant electrical energy.

[0025] The heating system according to the invention is all the more interesting in the presence of a plenum forming a passive air inlet, namely devoid of means of action for controlling the elements penetrating into the plenum, such as for example a particle filter.

[0026] According to a possibility compatible with the previous ones, the supply fluid connection and the discharge fluid connection can pass through the same wall of the plenum to reach an engine compartment of the aircraft, the turbine engine being at least partly housed in this engine compartment.

[0027] The arrangement of the heating system then has a reduced impact on the plenum.

[0028] According to a possibility compatible with the previous ones, the discharge fluid connection can be open to the engine compartment, the turbine engine being at least partly housed in this engine compartment.

[0029] The hot air passing through the heat exchanger is then ejected into the engine compartment.

[0030] Such a discharge has no impact on the aircraft insofar as the casing of the turbine engine reaches temperatures higher than the temperature reached by this air at the outlet of the heat exchanger. By way of illustration, using the aforementioned P3 port, the hot air at the outlet of the discharge fluid connection can be of the order of 100 to 120 degrees Celsius, while certain zones of the turbine engine can reach temperatures higher than 150 degrees Celsius. In addition, the engine compartment can comprise a fire detector which generates an alert signal in the presence of a temperature higher than 200 degrees Celsius in the engine compartment, and therefore in the presence of a temperature higher than the temperature of the hot air leaving the heat exchanger.

[0031] Usually, the engine compartment can open onto an outlet nozzle, which sucks in the air present in this engine compartment.

[0032] According to a possibility compatible with the previous ones, the heat exchanger can comprise two walls separated by pads, said hot air circulating between the two walls.

[0033] The heat exchanger may be relatively flat, and may have limited impact on the operation of the air inlet. The two walls can be parallel to each other. For example, the walls each have a substantially parallelepiped shape when viewed from above, the heat exchanger can have a block shape.

[0034] According to a possibility compatible with the previous one, the heat exchanger may comprise a central deflector providing a U-shaped path between the two walls, this path going from the supply fluid connection to the discharge fluid connection according to a direction of circulation of the hot air.

[0035] This feature makes it possible to arrange the supply fluid connection and the discharge fluid connection on the same side of the heat exchanger, with respect to a direction of extension of the heat exchanger. Therefore, the arrangement of the heating system within an aircraft can be facilitated.

[0036] According to a possibility compatible with the previous ones, the heating system may comprise one or more fixings attaching the heat exchanger to the plenum, the heat exchanger not being in contact with at least one bottom of the plenum.

[0037] For example, four fasteners are connected to four corner areas of the heat exchanger, possibly located under the heat exchanger or near the heat exchanger.

[0038] The impact on the plenum is thus limited. In addition, this arrangement makes it possible to generate an air stream surrounding the heat exchanger. This arrangement can make it possible to promote convection heating and / or to arrange drains under the heat exchanger to evacuate the water.

[0039] According to a possibility compatible with the previous ones, the fluid supply connection may comprise a solenoid valve arranged between two pipes, the solenoid valve being configured to authorize or prohibit the routing of said hot air into the heat exchanger.

[0040] A human-machine interface may be connected by a wired or wireless connection to the solenoid valve to allow a pilot to control this solenoid valve. Thus, a pilot can open the solenoid valve only when the aircraft is operating in icing conditions.

[0041] According to a possibility compatible with the previous ones, the fluid supply connection may include a restriction forming a flow limiter.

[0042] The restriction can make it possible to maintain a movement of hot air in the heat exchanger at a subsonic speed, while having an optimal fluid flow rate in order to obtain an optimized heat exchange. The heating system can thus be relatively simple and easily certifiable by the aeronautical authorities.

[0043] According to a possibility compatible with the previous ones, the heating system can include a pressure sensor connected to an alerter.

[0044] The pressure sensor is a sensor for generating a signal in the presence of a pressure greater than a threshold in the supply fluid connection. The pressure sensor simply has the function of indicating to a crew whether the heating system is working or not.

[0045] If the heating system is not operating, the pressure in the supply fluid connection is below the threshold. A pilot is warned in order to escape icing conditions as quickly as possible.

[0046] For example, the pressure sensor includes a pressure switch that emits a signal when the pressure in the supply fluid connection is greater than or equal to the threshold. The term "signal" may refer to an analog or digital, electrical or optical signal. The alerter may be configured to generate an alert until the signal is received.

[0047] According to a possibility compatible with the previous ones, the plenum may comprise at least one drain.

[0048] Such a drain may comprise a simple hole or a more complex device comprising for example a valve and / or a grid.

[0049] At least one drain may be located on a wall of the plenum opposite the heat exchanger.

[0050] For example, the plenum may include four drains placed in four corners to drain the plenum regardless of the roll and pitch angles of the aircraft.

[0051] According to a possibility compatible with the previous ones, the air supply system can be a passive system ingesting air coming from said external environment under the effect of suction from the turbine engine.

[0052] This system is in fact distinct from a dynamic system subject to other phenomena.

[0053] The invention and its advantages will appear in more detail in the context of the description which follows with examples given for illustrative purposes with reference to the appended figures which represent:

[0054] [Fig.l], an exterior view of an aircraft according to the invention,

[0055] [Fig.2], a sectional diagram of a plenum of an aircraft according to the invention,

[0056] [Fig. 3], a diagram illustrating the heating system of the invention, and

[0057] [Fig.4] a three-dimensional view of a heating system of the invention.

[0058] Elements present in several distinct figures are assigned a single reference.

[0059] [Fig.l] shows an aircraft 1 according to the invention. This aircraft 1 comprises a cell 2 extending in the direction of advance 500 of the aircraft from a rear end 4 towards a nose 3.

[0060] The aircraft 1 comprises a power plant equipped with a fuel supply system air with radial inlet provided with a plenum 40 with external radial opening 41 to feed a turbine engine. The term "radial" refers to a direction orthogonal to a central axis along which the turbine engine extends. For example, the turbine engine sets in motion a power transmission chain 35, this power transmission chain 35 being able to set in motion at least one rotor 5, 6 participating in the propulsion and / or the lift and / or the control of this aircraft 1. For example, the aircraft 1 is a helicopter provided with a main rotor 5 participating in its lift and its propulsion, as well as a tail rotor 6 participating in the control of the yaw movement of the aircraft 1.

[0061] With reference to [Fig.2], the plenum 40 comprises a duct 46 provided with an external opening 41 open onto an environment EXT external to the aircraft, or even an internal opening 42 open onto an air intake of a turbine engine 10. The duct 46 can be described as annular insofar as it extends radially from the external opening 41 towards the internal opening 42. To delimit the duct 46, the plenum 40 can comprise two partitions 43, 44, invisible in [Fig.2], connected by an edge 45, for example substantially in an arc of a circle.

[0062] With reference to [Fig. 3], the turboshaft engine 10 is arranged at least partially in an engine compartment 9. This engine compartment 9 can open onto an outlet nozzle 90, the air present in the engine compartment 9 being sucked in operation into the outlet nozzle 90.

[0063] The turbine engine 10 comprises a gas generator 15. The gas generator 15 is provided with a compression assembly 20 supplied with fresh air by the plenum 40. For example, the compression assembly 20 comprises one or more compression stages 21, 23. The example given illustrates a compression assembly 20 provided with a first compression stage 21 integral in rotation with a second compression stage 23, via a shaft 22.

[0064] Downstream of the compression assembly 20 in the direction of circulation of the gases within the turbine engine 10, the gas generator 15 comprises a combustion chamber 24, then a turbine assembly 25. The turbine assembly 25 is set in motion by the gases leaving the combustion chamber 24, and is rotationally integral with the compression assembly 20. The turbine assembly 25 may comprise at least one turbine. Finally, the turbine engine 10 comprises at least one working turbine 30, for example connected to the power transmission chain 35 mentioned above.

[0065] The turbine engine 10 therefore comprises a gas stream 26 which starts from the plenum 40 and passes through the blades of the compression stages 22, 23, then through the combustion chamber 24, the blades of the turbine(s) of the turbine assembly 25 and finally the blades of the working turbine(s) 30.

[0066] The aircraft 1 comprises a heating system 50 to minimize the accumulation of ice and / or snow in the plenum 40. The heating system 50 may be a de-icing system to limit the formation of ice and snow to an acceptable level for the turbine engine 10.

[0067] This heating system 50 comprises a heat exchanger 60 arranged in the volume 47 delimited by the plenum 40, i.e. inside the plenum 40 and not inside a wall of the plenum 40.

[0068] For example, the heating system 50 comprises one or more fasteners 85 each attaching the heat exchanger 60 to the plenum 40. For example, the heat exchanger is fixed by two fasteners respectively to the two partitions 43, 44 and by two other fasteners to the edge 45.

[0069] Optionally, the heat exchanger 60 is attached to the plenum 40 so as not to be in contact with the plenum 40, and in particular by being separated by a clearance from at least one bottom 400 of the plenum 40 or also from the partitions 43, 44. Thus, a clearance 300, for example of the order of 8 to 10 millimeters, separates the heat exchanger 60 from the bottom 400 of the plenum 40 to allow the circulation of air under the heat exchanger 60, more precisely between the heat exchanger 60 and the plenum 40. The bottom of the plenum may comprise a part of the plenum located under the heat exchanger, in the absence of an overturning of the aircraft and / or for example when the aircraft is resting on flat ground.

[0070] In addition, the heating system 50 comprises a fluid supply connection 70 for fluidly connecting the heat exchanger 60 and the gas stream 26. For example, the fluid supply connection 70 is fluidly connected to the gas stream 26 downstream of a compression stage of the compression assembly 20, or even of the compression stage 23 located before the combustion chamber, in order to take hot and compressed air.

[0071] This fluid supply connection 70 may comprise one or more pipes 71, 72. The term “pipe” designates one or more pipes allowing the hot gas taken to circulate.

[0072] The fluid supply connection 70 may comprise a solenoid valve 75 connected by a first pipe 71 to the heat exchanger 60 and by a second pipe 72 to a pressure tap 73 of the turbine engine 10. The solenoid valve 75 may be controlled by a human-machine interface 750. The solenoid valve 75 may be a two-position valve making it possible either to prohibit the circulation of hot gas to the heat exchanger 60 or to authorize the circulation of hot gas to the heat exchanger 60.

[0073] The fluid supply connection 70 may comprise a pressure sensor 81, where appropriate downstream of the solenoid valve 75. The pressure sensor 81 is connected to an alerter 82, by a wired or wireless connection. The pressure sensor 81 may, for example for example, transmitting a signal, analog or digital, electrical or optical, to the alerter 82 when a pressure in the supply fluid connection 70 is greater than or equal to a threshold, or conversely when a pressure in the supply fluid connection 70 is less than the threshold. The alert generated may take the form of a visual alarm, for example by means of the emission of a light with a light-emitting diode or an equivalent or the display on a screen of one or more characters, an audible alarm, by means of a loudspeaker, and / or a haptic alarm, for example using a vibrating unit vibrating an organ held or worn by an individual.

[0074] The fluid supply connection 70 may comprise at least one restriction 76 forming a flow limiter. In the example illustrated, a restriction 76 is located downstream of the solenoid valve 75. Alternatively or additionally, a restriction 76 may be located upstream of the solenoid valve 75, for example on the pipe 72 to limit the speed of the air in the solenoid valve 75, or even upstream of the pipe 72.

[0075] Furthermore, the heating system 50 further comprises a discharge fluid connection 80 for discharging the hot air passing through the heat exchanger 60. For example, the discharge fluid connection 80 is open to the engine compartment 9.

[0076] With reference to [Fig.4], the supply fluid connection 70 and the discharge fluid connection 80 can pass through the same wall 44 of the plenum 40 to reach the engine compartment 9.

[0077] According to another aspect, the heating system 50 may comprise at least one support 94 connecting the supply fluid connection 70 or the discharge fluid connection 80 either to the plenum 40 or to a supporting structure (not shown) of the aircraft 1.

[0078] For this purpose, the heat exchanger 60 may comprise an internal space delimited by two walls 61, 62, and a peripheral edge 63 connecting the two walls 61, 62. These two walls 61, 62 may comprise an upper wall 61 and a lower wall 62 located under the upper wall 61, at least as long as the aircraft 1 rests on a substantially horizontal ground via a landing gear. The upper wall 61 and the lower wall 62 may be parallel to each other and / or substantially parallelepipedal in shape, or even identical. The supply fluid connection 70 and the discharge fluid connection 80 are each connected to one of the two walls 61, 62 to circulate the hot air in the internal space. For example, the supply fluid connection 70 and the discharge fluid connection 80 can be connected to the same wall, and more precisely to the upper wall 61 according to the example shown in [Fig.4].

[0079] According to another aspect, the heat exchanger 60 may comprise a diverter central 65, for example comprising a possibly straight partition arranged between the two walls 61, 62. The central deflector 65 provides a U-shaped path 66 for the hot air between the two walls 61, 62. This path 66 extends from the supply fluid connection 70 to the discharge fluid connection 80. Consequently, the supply fluid connection 70 and the discharge fluid connection 80 can be arranged on the same side of the heat exchanger 60 in a direction of length extension of this heat exchanger 60.

[0080] According to another aspect, the plenum 40 may comprise at least one drain 96. For example, at least one drain 96 is provided on a wall opposite the heat exchanger 60, or even located under the heat exchanger 60. For example, the plenum comprises four drains 96 located at four corners of the plenum.

[0081] Therefore, when the aircraft 1 flies in icing conditions, if necessary, a pilot operates the human-machine interface 750 to open the solenoid valve 75. The hot air circulating in the compression assembly 20 of the turbine engine 10 automatically circulates in the supply fluid connection 70. If necessary, the pressure sensor 81 detects a change in pressure and transmits a signal to the alerter 82 which issues an alert.

[0082] The hot air then flows into the heat exchanger 60, then is ejected into the engine compartment 9 by the discharge fluid connection 80. The walls 61, 62 of the heat exchanger 60 heat up and tend to heat the plenum 40. If necessary, the ice or snow present in the plenum 40 melts, and the water flows out of the plenum 40 through a drain 96.

[0083] Naturally, the present invention is subject to numerous variations as to its implementation. Although several embodiments have been described, it is understood that it is not conceivable to exhaustively identify all possible modes. It is of course conceivable to replace a means described by an equivalent means without departing from the scope of the present invention and the claims.

Claims

Claims

1. Aircraft (1) provided with a turbine engine (10) comprising a gas generator (15), the gas generator (15) comprising a compression assembly (20) supplying compressed air to a combustion chamber (24), the gas generator (15) comprising a turbine assembly (25) supplied with gas by the combustion chamber (24), the aircraft (1) having a radial inlet air supply system provided with a plenum (40) supplying air to the compression assembly (20), the plenum (40) comprising a duct (46) provided with an external opening (41) open to an external environment (EXT) located outside the aircraft (1), characterized in that said aircraft (1) comprises a heating system (50), the heating system (50) comprising a heat exchanger (60) arranged in a volume delimited by the plenum (40), the heating system (50) comprising a connection supply fluid (70) and a discharge fluid connection (80) connected to the heat exchanger (60),the fluid supply connection (70) conveying hot air from the turbine engine (10) into the heat exchanger (60).,

2. Aircraft according to claim 1, characterized in that the turbine engine (10) comprising a gas stream (26) extending from the plenum (40) passing successively through the compression assembly (20) then the combustion chamber (24) and the turbine assembly (25), the fluid supply connection (70) is in fluid connection with the gas stream (26) downstream of a compression stage (21) of the compression assembly (20).

3. Aircraft according to any one of claims 1 to 2, characterized in that the supply fluid connection (70) and the discharge fluid connection (80) pass through the same wall (44) of the plenum (40) to reach an engine compartment (9) of the aircraft (1), the turbine engine (10) being at least partly housed in this engine compartment (9).

4. Aircraft according to any one of claims 1 to 3, characterized in that the delivery fluid connection (80) is open onto an engine compartment (9) of the aircraft (1), the turbine engine (10) being at least partly housed in this engine compartment (9).

5. Aircraft according to any one of claims 1 to 4, characterized in that the heat exchanger (60) comprises two walls (61, 62) separated by pads (64), said hot air (95) circulating between the two walls (61, 62).

6. Aircraft according to claim 5, characterized in that the heat exchanger (60) has a central deflector (65) providing a U-shaped path (66) between the two walls (61, 62), this path (66) going from the supply fluid connection (70) to the discharge fluid connection (80) in a direction of circulation of the hot air.

7. Aircraft according to any one of claims 1 to 6, characterized in that the heating system (50) comprises one or more fasteners (85) attaching the heat exchanger (60) to the plenum (40), the heat exchanger (60) not being in contact with at least one bottom (400) of the plenum (40).

8. Aircraft according to any one of claims 1 to 7, characterized in that the fluid supply connection (70) comprises a solenoid valve (75) arranged between two pipes (71, 72), the solenoid valve (75) being configured to authorize or prohibit the routing of said hot air into the heat exchanger (60).

9. Aircraft according to any one of claims 1 to 8, characterized in that the fluid supply connection (70) comprises a restriction (76) forming a flow limiter.

10. Aircraft according to any one of claims 1 to 9, characterized in that the heating system (50) comprises a pressure sensor (81) connected to an alerter (82).

11. Aircraft according to any one of claims 1 to 10, characterized in that the plenum (40) comprises at least one drain (96).

12. Aircraft according to any one of claims 1 to 11, characterized in that the air supply system is a passive system ingesting air from said external environment (EXT) under the effect of suction from the turbine engine (10).

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