Frost protection system, aircraft engine nacelle having such system and aircraft
The heat transfer is fully covered in the aircraft engine nacelle through the heat exchanger device, which solves the problem of frost and ice interferes with the air supply, achieves an effective anti-frost effect and reduces costs.
Patent Information
- Application Number
- CN201811199379.8
- Authority / Receiving Office
- CN · China
- Patent Type
- Patents(China)
- Current Assignee / Owner
- Priority Date
- 2017-10-20
- Filing Date
- 2018-10-16
- Publication Date
- 2024-12-27
- Estimated Expiration
- 2038-10-16
AI Technical Summary
The air inlets of the aircraft engine nacelle are prone to form frost ice, which interferes with the air supply. The existing anti-frost system is insufficient in the transmission distance of hot air, resulting in the aerodynamic sensitivity of the sound absorption channel and high cost.
The heat exchanger device is adopted, including a heat pipe, an evaporator, a condenser and a heat transfer fluid, and heat transfer fluid is transferred to the sound absorption panel through a heat source. The evaporator extracts the heat from the hot air and transfers it to the condenser. The condenser heats the sound absorption panel to achieve Full coverage to prevent frost.
Effectively and inexpensively protect the sound-absorbing panel from frost damage, improves the anti-frost effect and enhances the aerodynamic stability of the air intake.
Smart Images
Figure CN109693795B_ABST
Abstract
Description
Technical Field
[0001] The invention relates to a frost protection system for an air inlet of an aircraft engine nacelle and to a nacelle and an aircraft provided with such a protection system. Background Art
[0002] The leading edge of an aircraft, in particular the air intake lip of an aircraft engine nacelle, may be subject to the formation of frost which accumulates to form ice blocks. The formation of these ice blocks may interfere with the air supply to the engine. For example, the ice blocks may become loose and collide with the engine fan blades. As a result, the fan blades tend to become brittle or even damaged thereby.
[0003] There are frost protection systems that extract hot air from the compression stage of the aircraft engine in order to inject it into an annular space located behind the lip of the nacelle. The hot air then flows through the annular space, heats the lip, and is conveyed into channels in the acoustic panels in order to heat the skin of said acoustic panels. However, the skin of the acoustic panels is heated over a short distance, which causes defrosting over a short distance. For very short air inlets, this distance may be insufficient. In particular, a shortening of the air inlet may cause the surface of the acoustic channel to become more aerodynamically sensitive. Summary of the invention
[0004] The object of the present invention is to overcome these drawbacks by providing an anti-frost protection system for a nacelle.
[0005] To this end, the invention relates to a frost protection system for an aircraft engine nacelle, the nacelle comprising an inner shroud provided with at least one sound absorbing panel.
[0006] According to the invention, the protection system comprises a heat exchanger device comprising at least one heat pipe configured for transferring heat emitted by a heat source to the one or more sound absorbing panels.
[0007] Thus, thanks to the invention, the acoustic panels are more effectively and cost-effectively protected from frost damage by virtue of the one or more heat pipes.The heat emitted by the heat source is used to defrost all the acoustic panels of the nacelle, not just a portion of the acoustic panels located near the lip.
[0008] According to one feature, the heat exchanger device further comprises:
[0009] - heat transfer fluid;
[0010] - at least one evaporator, said at least one evaporator being thermally connected to said heat source, said one or more evaporators being configured for extracting at least a portion of the heat supplied by said heat source, said extracted heat being transferred to said heat transfer fluid;
[0011] - at least one condenser attached to the inner shroud, the one or more condensers being configured for delivering at least a portion of the heat extracted by the one or more evaporators to the one or more acoustic panels, the extracted heat being transferred to the one or more condensers via the heat transfer fluid;
[0012] Each of the evaporators is fluidly connected to at least one condenser via at least one heat pipe through which the heat transfer fluid flows.
[0013] Additionally, the one or more heat pipes include:
[0014] at least one steam pipe configured to convey the heat transfer fluid vaporized by the heat extracted by the evaporator from the evaporator to the condenser,
[0015] - at least one liquid pipe configured to transport the heat transfer fluid condensed by cooling in the condenser from the condenser to the evaporator.
[0016] Furthermore, the one or more condensers comprise one or more heating channels.
[0017] The invention also relates to an aircraft engine nacelle comprising an inner shroud provided with at least one sound absorbing panel.
[0018] According to the invention, the nacelle comprises a frost protection system as described above.
[0019] In addition, the nacelle comprises an air inlet lip forming a leading edge of the nacelle, the lip having an annular space, the annular space being enclosed by an internal baffle and being arranged for receiving a supply of hot air, the one or more evaporators being connected to the internal baffle by attachment, the one or more evaporators being configured for extracting at least a portion of the heat supplied by the hot air supplied to the annular space of the lip through the internal baffle, the extracted heat being transferred to the heat transfer fluid.
[0020] The invention also relates to an aircraft, in particular a transport aircraft, equipped with at least one engine surrounded by a nacelle comprising an inner shroud provided with at least one sound absorbing panel.
[0021] According to the invention, the aircraft comprises a frost protection system as described above.
[0022] According to one embodiment, the nacelle comprises an air inlet lip forming a leading edge of the nacelle, the lip having an annular space closed by an internal baffle and arranged for receiving a supply of hot air,
[0023] The heat source corresponds to the internal partition,
[0024] the one or more evaporators being connected to the inner baffle by attachment, the one or more evaporators being configured to extract at least a portion of the heat supplied by the hot air supplied to the annular space of the lip through the inner baffle, the extracted heat being transferred to the heat transfer fluid,
[0025] The aircraft comprises at least one air heating device configured for generating the hot air supplied to the annular space of each of the nacelles.
[0026] In addition, the aircraft comprises:
[0027] - at least one duct connecting the one or more air heating devices to the annular space of each of the nacelles, the one or more ducts being configured for conveying the hot air generated by the air heating devices to the annular space of the lip,
[0028] - at least one valve for each of said ducts, said at least one valve being configured for regulating the pressure and flow rate of said hot air flowing through said one or more ducts.
[0029] Furthermore, the air heating device corresponds to a compression stage of the engine enclosed by the nacelle.
[0030] According to another embodiment, the aircraft comprises an electrical system corresponding to the heat source.
[0031] According to a variant embodiment, the electrical system corresponds to an electrical device dedicated to generating heat for the frost protection system.
[0032] According to another variant, the electrical system corresponds to a typical electrical device dedicated to powering the aircraft. BRIEF DESCRIPTION OF THE DRAWINGS
[0033] The present invention together with its features and advantages will become more apparent after reading the description given with reference to the accompanying drawings, in which:
[0034] - Figure 1 shows a side view of an aircraft engine including a hot air supply to the annular space of the lip,
[0035] - Figure 2 shows a longitudinal section of a lip comprising an anti-frost protection system according to one embodiment,
[0036] - Figure 3 shows a cross-sectional view of a lip including a frost protection system according to one embodiment,
[0037] - Figure 4 shows a schematic diagram of an anti-frost protection system according to an embodiment,
[0038] - Figure 5 shows a schematic diagram of a condenser according to an embodiment,
[0039] - Figure 6 A side view of an aircraft including a frost protection system according to one embodiment is shown. DETAILED DESCRIPTION
[0040] Figure 2 Schematically shows a nacelle 2 for an aircraft AC engine 3 ( Figure 6 ) is an embodiment of an anti-frost protection system.
[0041] Engine 3 Nacelle 2 refers to the space surrounding the aircraft AC engine 3 ( Figure 1 ) (e.g. an aircraft AC turbofan engine).
[0042] The nacelle generally comprises an outer shroud 4, an inner shroud 5, and an air inlet lip 7. The two shrouds 4 and 5 are generally coaxial and form a space between them. The lip 7 connects the two shrouds 4 and 5 together.
[0043] The outer shroud 4 forms the outer covering of the nacelle 2. The inner shroud 5 is provided with at least one sound absorbing panel 6. The lip 7 forms the leading edge of the nacelle 2.
[0044] The nacelle is often equipped with a sound absorbing panel covering the inner wall of the nacelle 2 at the air inlet upstream of the fan 22. Generally, the sound absorbing panel 6 has a sandwich structure comprising one or more honeycomb honeycomb structure layers suitable for capturing noise. This honeycomb structure layer has an outer face covered by a porous layer (called sound absorbing skin) and an inner face covered by an impermeable layer (called dense skin).
[0045] The protection system 1 comprises a heat exchanger arrangement 11 configured for transferring heat 14 emitted by a heat source to the one or more acoustic absorbing panels 6 .
[0046] The heat exchanger arrangement 11 comprises at least one heat pipe 12 configured for transferring heat 14 from a heat source to the one or more acoustic absorbing panels 6 .
[0047] Heat pipe 12 generally refers to a heat conducting element that operates on the principle of heat transfer through phase change of a fluid.
[0048] The heat exchanger device 1 further comprises a heat transfer fluid and at least one evaporator 13 thermally connected to the heat source. The one or more evaporators 13 are configured to extract at least a portion of the heat 14 supplied by the heat source. The heat 14 is then transferred to the heat transfer fluid ( Figure 2 , Figure 3 ,and Figure 4 ).
[0049] The heat exchanger device 1 further comprises at least one condenser 15 attached to the inner shroud 5. The heat 14 extracted by the one or more evaporators 13 is transferred to the one or more condensers 15 via a heat transfer fluid. The one or more condensers 15 are configured to deliver at least a portion of the heat 14 extracted by the one or more evaporators 13 to the one or more sound absorbing panels 6.
[0050] Preferably, the condenser 15 is spread over the entire inner shroud 5 .
[0051] Each evaporator 13 is connected by at least one heat pipe 12 ( Figure 2 , Figure 3 ,and Figure 4 ) is fluidly connected to at least one condenser 15.
[0052] Advantageously, the one or more heat pipes 12 include at least one steam pipe 16, which is configured to transport the heat transfer fluid vaporized by the heat 14 extracted by the evaporator 13 from the evaporator 13 to the condenser 15. The one or more heat pipes 12 also include at least one liquid pipe 17, which is configured to transport the heat transfer fluid condensed by cooling in the condenser 15 from the condenser 15 to the evaporator 13.
[0053] The liquid pipe 17 may be a pipe that allows the condensed heat transfer fluid to return to the evaporator 13 by gravity or capillary action.
[0054] According to one embodiment, the one or more heat pipes 12 include a central pipe and a peripheral pipe surrounding the central pipe. The central pipe may correspond to the steam pipe 16, and the peripheral pipe may correspond to the liquid pipe 17.
[0055] Advantageously, the one or more condensers 15 comprise a plurality of acoustic panels 6 ( Figure 5 ) in one or more heating channels 18.
[0056] The lip 7 has an annular space 8 which is closed by an internal partition 9. The internal partition 9 separates the annular space 8 from the rest of the space formed between the two shrouds 4 and 5.
[0057] Generally, the annular space 8 is constituted by two D-shaped ducts forming a ring just behind the leading edge.
[0058] The annular space 8 of the lip is arranged for receiving a supply 10 of hot air.
[0059] According to one embodiment, the heat source corresponds to the internal partition 9 heated by hot air 10 supplied to the annular space 8 of the lip 7 .
[0060] Without limitation, the internal baffles 9 are typically heated to a temperature ranging from 250°C to 450°C.
[0061] Advantageously, at least one heat pipe 12 is configured for transferring heat 14 from the internal partition 9 to the one or more acoustic absorbing panels 6 .
[0062] The one or more evaporators 13 are thus configured to extract at least a portion of the heat 14 supplied by the hot air 10 supplied to the annular space 8 of the lip 7 through the internal baffle 9. The heat 14 is then transferred to the heat transfer fluid ( Figure 2 , Figure 3 ,and Figure 4 ).
[0063] Preferably, the evaporator 13 is spread over the entire inner partition 9 .
[0064] The hot air 10 supplied to the annular space 8 may be generated by an air heating device 19 of the aircraft AC. The air heating device 19 is configured for generating the hot air 10 supplied to the annular space 8 of each nacelle 2.
[0065] For example, the aircraft AC comprises at least one duct 20, which couples the one or more air heating devices 19 to the annular space 8 of each nacelle 2. The one or more ducts 20 are configured to deliver the hot air 10 generated by the air heating device 19 to the annular space 8 of the lip 7. The aircraft AC also comprises at least one valve 21 for each duct 20, which is configured to adjust the pressure and flow rate of the hot air 10 flowing through the one or more ducts 20.
[0066] The one or more tubes 20 may correspond to a pipe or a piccolo tube.
[0067] For example, the air heating device 19 corresponds to a compression stage of the engine 3 enclosed by the nacelle 2. Thus, the compression stage of the engine 3 supplies hot air 10 to the annular space 8 of the lip 7 of the nacelle 2 enclosing the engine 3.
[0068] Thus, the heating device 19 supplies hot air 10 to the annular space 8 of the lip 7. The hot air 10 then flows through the annular space 8 of the lip 7 and heats the internal baffles 9. Then, the one or more evaporators 13 attached to the internal baffles 9 extract the heat of the internal baffles 9 heated by the hot air 10. To this end, the heat transfer fluid in the one or more evaporators 13 is vaporized and transported to the one or more condensers 15 through the steam pipes 16 of the one or more heat pipes 12. Heat is thus transferred to the sound absorbing panel 6 via the one or more condensers 15, wherein the heat transfer fluid is condensed by supplying heat to the condenser 15. The condensed heat transfer fluid is then returned to the one or more evaporators 13 via the one or more heat pipes 12.
[0069] This protection system 1 allows efficient transfer of heat from the interior partition 9 to the acoustic panel 6. In addition, the protection system 1 makes it possible to use the heat at the interior partition 9 heated by the hot air and therefore allows more efficient use of the heat supplied by the hot air 10 generated by the heating device 19.
[0070] According to another embodiment, the heat source corresponds to the electrical system of the aircraft AC (not shown).
[0071] According to a variant of this embodiment, the electrical system corresponds to an electrical device dedicated to generating heat for the protection system 1. This electrical device can be powered by the electrical core of the aircraft AC or directly by the generators of the engines.
[0072] According to another variant, the electrical system corresponds to typical electrical devices dedicated to supplying AC power to an aircraft. For example, in order for the protection device 1 to exploit the excess heat, the electrical devices are heated to a greater degree than is necessary.
[0073] By virtue of this embodiment, it is possible to avoid integrating electrical systems and power supply cables in the one or more sound absorbing panels 6 .
Claims
1. A nacelle for an aircraft engine, comprising: An inner shroud (5) is provided with at least one sound absorbing panel (6) and Frost protection system (1), The frost protection system (1) comprises a heat exchanger device (11), the heat exchanger device comprising at least one heat pipe (12), the at least one heat pipe being configured to transfer heat (14) emitted by a heat source to the at least one sound absorbing panel (6); The heat exchanger device (11) further comprises: - heat transfer fluid, at least one evaporator (13) thermally connected to the heat source, the at least one evaporator (13) being configured to extract at least a portion of the heat (14) supplied by the heat source, the extracted heat (14) being transferred to the heat transfer fluid; at least one condenser (15) attached to the inner shroud (5), the at least one condenser (15) being configured to deliver at least a portion of the heat (14) extracted by the at least one evaporator (13) to the at least one acoustic panel (6), the extracted heat (14) being transferred to the at least one condenser (15) via the heat transfer fluid; Each of the evaporators (13) is fluidly connected to at least one condenser (15) via at least one heat pipe (12), the heat transfer fluid flowing through the at least one heat pipe, and The nacelle (2) comprises an air inlet lip (7) forming a leading edge of the nacelle (2), the air inlet lip (7) having an annular space (8) which is closed by an internal baffle (9) and is arranged to receive a supply of hot air, characterised in that the heat source corresponds to the internal partition (9), and The at least one evaporator (13) is connected to the internal baffle (9) by attachment, and the at least one evaporator (13) is configured to extract at least a portion of the heat (14) supplied by the hot air supplied to the annular space (8) of the air inlet lip (7) through the internal baffle (9), the extracted heat (14) being transferred to the heat transfer fluid.
2. The nacelle for an aircraft engine according to claim 1, It is characterized in that The at least one heat pipe (12) comprises: at least one steam pipe (16) configured to convey the heat transfer fluid vaporized by the heat (14) extracted by the evaporator (13) from the evaporator (13) to the condenser (15), - at least one liquid pipe (17), the at least one liquid pipe being configured to transport the heat transfer fluid condensed by cooling in the condenser (15) from the condenser (15) to the evaporator (13).
3. The nacelle for an aircraft engine according to claim 2, It is characterized in that The at least one condenser (15) comprises one or more heating channels (18).
4. An aircraft equipped with at least one engine (3), It is characterized in that The aircraft comprises a nacelle according to any one of claims 1 to 3, enclosing the at least one engine (3).
5. The aircraft according to claim 4, It is characterized in that The aircraft (AC) comprises: at least one air heating device (19) configured to generate hot air supplied to the annular space (8) of each of the nacelles (2), - at least one duct (20) connecting the at least one air heating device (19) to the annular space (8) of each of the nacelles (2), the at least one duct (20) being configured to convey the hot air generated by the air heating device (19) to the annular space (8) of the air inlet lip (7), - at least one valve (21) for each of said ducts (20), said at least one valve being configured for regulating the pressure and flow rate of said hot air (10) flowing through said at least one duct (20).
6. The aircraft according to claim 5, It is characterized in that The air heating device (19) corresponds to a compression stage of the engine (3) enclosed by the nacelle (2).