Universal test flight platform of turboshaft engine

Abstract

The invention relates to the technical field of aircraft trial, and provides a universal test flight platform of a turboshaft engine, which comprises an aircraft, a first test cabin, a first power pack, an independent reduction gearbox and a propeller, two wings are respectively arranged on two sides of the aircraft, the first test cabin is arranged on one wing, the first test cabin and the first power pack are positioned on the same side of the aircraft, and the independent reduction gearbox is arranged on the first test cabin. The first power pack is arranged at the joint of a wing and the aircraft, the first test chamber is provided with a containing cavity, a mounting part is arranged in the containing cavity and used for mounting an engine and enabling the axial direction of the engine to be consistent with that of the containing cavity, the engine is connected with the first power pack through the first test chamber, and the independent reduction gearbox is arranged in the containing cavity. The propeller is arranged outside the containing cavity and located at the end of the first test chamber, and the engine is in transmission connection with the independent reduction gearbox and the propeller in sequence. The turboshaft engine is well matched with the test flight platform, test flight verification of the whole flight envelope of the turboshaft engine is met, and the test flight platform has the characteristic of universality.

Description

Technical Field

[0001] The invention belongs to the technical field of engine flight test, and in particular relates to a universal flight test platform for a turboshaft engine. Background Art

[0002] The existing turboshaft engine flight test platforms are mainly helicopter platforms. Since there are many requirements for the adaptation between helicopter platforms and turboshaft engines in terms of mechanical structure layout, water and electricity supply interface requirements, flight engine performance docking and speed matching, it is difficult and costly. In addition, the flight speed envelope of the helicopter platform is narrower than that of the turboshaft engine, and it cannot meet the verification requirements for the turboshaft engine speed envelope. The existing general flight test platforms are usually fixed-wing flight platforms, which generally cannot meet the requirements for the output power absorption of turboshaft engines, and the functions of the two are not completely matched. Therefore, using the helicopter platform as a test flight platform for the turboshaft engine requires adaptive modification, which has many requirements, great difficulty and high cost, resulting in an extension of the development cycle of the turboshaft engine and inconvenience in conducting flight tests. Summary of the invention

[0003] In view of the above problems, the present invention proposes a universal flight test platform for a turboshaft engine, comprising: an aircraft, a first test cabin, a first power pack, an independent reduction gearbox and a propeller, two wings are respectively provided on both sides of the aircraft, the first test cabin is arranged on one of the wings, the first test cabin and the first power pack are located on the same side of the aircraft, the first power pack is arranged at the connection between one of the wings and the aircraft, the first test cabin has a accommodating cavity, a mounting portion is provided in the accommodating cavity, the mounting portion is used to install the engine and make the axial direction of the engine consistent with that of the accommodating cavity, the engine is connected to the first power pack through the first test cabin, the independent reduction gearbox is arranged in the accommodating cavity, the propeller is arranged outside the accommodating cavity and located at the end of the first test cabin, and the engine is sequentially connected to the independent reduction gearbox and the propeller.

[0004] Optionally, an air intake duct is provided in the accommodating cavity, one end of the air intake duct passes through the accommodating cavity and communicates with the outside, and the other end is communicated with a flow channel of the engine, and a part of the air intake duct is between the engine and the independent reduction box.

[0005] Optionally, a distortion test section is provided between the air intake duct and a flow passage of the engine.

[0006] Optionally, a water spraying and freezing test section is further provided in the accommodating chamber, and the water spraying and freezing test section is connected to the air inlet duct through a pipeline.

[0007] Optionally, the first test chamber includes an outer shell and an exhaust duct, the outer shell circumferentially surrounds the accommodating cavity, the exhaust duct and the propeller are respectively located at two ends of the outer shell, and the exhaust duct is connected to the outer shell so that the flow channel in the exhaust duct is connected to the accommodating cavity.

[0008] Optionally, the mounting portion includes a main fulcrum mounting frame and a secondary fulcrum mounting frame, the engine is mounted on the main fulcrum mounting frame and the secondary fulcrum mounting frame, and the main fulcrum mounting frame and the secondary fulcrum mounting frame are arranged on the cavity wall of the accommodating cavity at a predetermined interval.

[0009] Optionally, a distortion plate is provided at the connection between the exhaust duct and the engine.

[0010] Optionally, a power interface and a gas circuit interface are provided on the first test chamber, and the power interface or the gas circuit interface cooperates with a corresponding structural interface on the first power pack.

[0011] Optionally, a live-fire shooting device is provided outside the first test chamber.

[0012] Optionally, it also includes a second test chamber and a second power pack, wherein the second test chamber and the second power pack are located on the same side of the aircraft, and the second test chamber is symmetrical to the first test chamber about the center line of the aircraft, and the second power pack is symmetrical to the first power pack about the center line of the aircraft.

[0013] Compared with the prior art, the present invention has the following advantages:

[0014] The engine is reasonably arranged in the accommodating cavity of the first test cabin through the installation part, and the engine is connected to the first power pack through the first test cabin, the independent reduction box is arranged in the accommodating cavity, and the propeller is arranged outside the accommodating cavity and located at the end of the first test cabin. The independent reduction box and the propeller are sequentially connected with the engine in a transmission manner, so as to achieve a good adaptation of the turboshaft engine and the test flight platform, which can meet the test flight verification of the full flight envelope of the turboshaft engine, and has the ability to carry out test flights of various types of turboshaft engines, so that this platform has universal characteristics. In addition, the propeller pitch of the propeller is adjusted to realize the power conversion of the turboshaft engine, and the transmission ratio of the independent reduction box is adjusted to realize the power absorption matching the propeller and the turboshaft engine.

[0015] Other features and advantages of the present invention will be described in the following description, and partly become obvious from the description, or be understood by implementing the present invention. The purpose and other advantages of the present invention can be realized and obtained by the structures pointed out in the description and the drawings. BRIEF DESCRIPTION OF THE DRAWINGS

[0016] In order to more clearly illustrate the embodiments of the present invention or the technical solutions in the prior art, the following briefly introduces the drawings required for use in the embodiments or the description of the prior art. Obviously, the drawings described below are some embodiments of the present invention. For ordinary technicians in this field, other drawings can be obtained based on these drawings without paying creative work.

[0017] Figure 1 A schematic diagram of a general flight test platform for a turboshaft engine in an embodiment of the present invention is shown;

[0018] Figure 2 A schematic diagram showing the layout of a test cabin of a general test flight platform for a turboshaft engine in an embodiment of the present invention is shown;

[0019] Figure 3 A control schematic diagram of a universal flight test platform for a turboshaft engine in an embodiment of the present invention is shown.

[0020] In the figure, 1. aircraft; 2. first test cabin; 20. accommodating chamber; 21. outer shell; 22. exhaust duct; 23. air inlet duct; 24. distortion test section; 25. water spray icing test section; 3. first power pack; 4. independent reduction gearbox; 5. propeller; 6. engine; 7. mounting section; 71. main fulcrum mounting frame; 72. secondary fulcrum mounting frame; 8. second test cabin; 9. second power pack. DETAILED DESCRIPTION

[0021] In order to make the purpose, technical solution and advantages of the embodiments of the present invention clearer, the technical solution in the embodiments of the present invention will be clearly and completely described below in conjunction with the drawings in the embodiments of the present invention. Obviously, the described embodiments are part of the embodiments of the present invention, not all of the embodiments. Based on the embodiments of the present invention, all other embodiments obtained by ordinary technicians in this field without creative work are within the scope of protection of the present invention.

[0022] In the general flight test platform of the engine, new engine models often adopt a large number of new designs, new processes and new materials. In the development of new engine models, after completing the necessary ground tests, they should be installed for flight tests as soon as possible to verify the performance and functions of the engine, expose the engine defects as early as possible, and provide first-hand data and information for design modification and the next development decision, thereby shortening the engine development cycle and reducing development risks.

[0023] Although the simulation capability of aircraft / engines has been significantly enhanced, and the test verification capability of components and complete machines has also been greatly improved, the utilization rate of flight platforms by foreign engine manufacturers has not decreased, but has shown an upward trend. The purpose of engine flight tests is to preliminarily verify the performance and function of the engine, expose existing faults and design defects, and provide a decision-making basis for modifying the design and the next step of development. According to the development experience of Chinese and foreign engines, in order to reduce the risk of prototype testing, a type of engine must undergo flight tests before the formal finalization / identification flight test. In addition, for the research and development of major technical problems encountered during the finalization flight test, if the prototype is a single-engine platform, in order to resolve the test flight risk, a flight platform can also be selected to carry out research and development test flights. However, the existing general platform cannot adapt to turboshaft engines, which brings inconvenience to flight tests.

[0024] like Figure 1 As shown, the present invention provides a universal test flight platform for a turboshaft engine, see Figure 2 , including: an aircraft 1, a first test cabin 2, a first power pack 3, an independent reduction box 4 and a propeller 5. Two wings are respectively provided on both sides of the aircraft 1, and the first test cabin 2 is arranged on one wing. By installing the first test cabin 2 on the wing of the fixed-wing aircraft 1, it is possible to verify the dual-engine matching and fault handling countermeasures of the turboshaft engine 6 on the test flight platform. The first power pack 3 is arranged at the connection between the wing and the aircraft 1. The first test cabin 2 and the first power pack 3 are located on the same side of the aircraft 1. The first test cabin 2 has a accommodating cavity 20, and a mounting portion 7 is provided in the accommodating cavity 20. The mounting portion 7 is used to install the engine 6 and make the axial direction of the engine 6 consistent with that of the accommodating cavity 20. The first test cabin 2 realizes the installation structure layout of the simulated turboshaft engine 6 on the helicopter. The engine 6 is connected to the first power pack 3 through the first test chamber 2, the independent reduction box 4 is arranged in the accommodating chamber 20, the propeller 5 is arranged outside the accommodating chamber 20 and located at the end of the first test chamber 2, the engine 6 is sequentially connected to the independent reduction box 4 and the propeller 5, and the propeller 5 is detachably connected to the engine 6. Therefore, by replacing other propellers 5, the propeller diameter and the number of blades of other propellers 5 can be changed, and the power conversion of the turboshaft engine 6 can be realized by controlling and adjusting the adjustable pitch of the propeller 5. The transmission ratio of the independent reduction box 4 is adjusted to achieve the matching power absorption of the turboshaft engine 6. By reasonably arranging the engine 6 and other related components, the turboshaft engine 6 can be well adapted to the test flight platform, which can meet the test flight verification of the full flight envelope of the turboshaft engine 6, and has the ability to carry out multiple types of turboshaft engine 6 test flights, so that this platform has universal characteristics.

[0025] It should be noted that the first power pack 3 not only serves as the main power source of the fixed-wing aircraft 1 to ensure that the fixed-wing aircraft 1 can reach the flight state required for the test flight conditions of the turboshaft engine 6, but also can serve as the source of electricity and airflow required by the first test cabin 2 to meet the test requirements of the first test cabin 2 and optimize the utilization of the capabilities of the general test flight platform.

[0026] The independent reduction box 4 connects the turboshaft engine 6 with the propeller 5, and the output shaft speed of the turboshaft engine 6 is matched with the speed of the propeller 5 by adjusting the reduction ratio of the independent reduction box 4, so as to meet the requirement that the first test chamber 2 can match the output shaft speeds of different turboshaft engines 6. By adjusting the interface between the independent reduction box 4 and the turboshaft engine 6, the test flight requirements of turboshaft engines 6 of different power levels can be met.

[0027] The propeller 5 is designable and changeable, and is also mechanically adjustable. The absorbed power of the propeller 5 can be adjusted by the propeller pitch, propeller diameter and number of blades. Since the turboshaft engine 6 adopts a control method of constant output shaft speed, the propeller 5 that meets the power requirement is designed and changed according to the output power of the turboshaft engine 6, and the propeller pitch of the propeller 5 is mechanically adjustable to adapt to the power load change of the turboshaft engine 6.

[0028] The aircraft 1 is capable of meeting the flight envelope required for the test flight of the turboshaft engine 6; it can be set with a variety of turboshaft engine 6 flight fault simulation functions, such as air restart, windmill start, control system and other fault simulations. The first power pack 3 provides flight power for the test flight platform and reaches the flight conditions for the test, and then sends instructions through the control system of the first test cabin 2 to complete the test flight test of the turboshaft engine 6 flight fault simulation.

[0029] In one embodiment, an air inlet 23 is provided in the accommodating cavity 20, one end of the air inlet 23 passes through the accommodating cavity 20 and communicates with the outside, and the other end communicates with the flow channel of the engine 6, and a part of the air inlet 23 is between the engine 6 and the independent reduction box 4. It should be noted that the outer shape structure of the air inlet 23 is an L-shaped structure, so that a part of the air inlet 23 is arranged in the radial direction of the accommodating cavity 20 and communicates with the flow channel of the engine 6, and the other part is arranged in the circumferential direction of the accommodating cavity 20 and passes through the cavity to communicate with the outside. Such an arrangement of the air inlet 23 can meet the installation of turboshaft engines 6 with different sizes of air inlet 23, so as to facilitate the replacement of turboshaft engines 6 of different types and corresponding air inlets 23 at any time.

[0030] In one embodiment, a distortion test section 24 is provided between the air inlet 23 and the flow passage of the engine 6. The distortion test section 24 is used to perform a distortion test on the air inlet 23 of the engine 6, and the distortion test section 24 is arranged between the air inlet 23 and the flow passage of the engine 6, so as to be more suitable for the turboshaft engine 6, and to perform a simulated intake distortion test, including a distortion test on the intake temperature.

[0031] In one embodiment, a water spray icing test part 25 is further provided in the accommodating chamber 20, and the water spray icing test part 25 is connected to the air inlet 23 through a pipe. Specifically, the water spray icing test part 25 is arranged along the axial direction of the accommodating chamber 20, and the water spray icing test part 25 is close to the air inlet 23. The water spray icing test part 25 can spray water into the air inlet 23, and the water flow temperature is based on the environmental icing test of the simulated turboshaft engine 6. When the flight platform flies in a high-altitude low-temperature environment, the water spray icing test part 25 is turned on to simulate the ice swallowing ability of the turboshaft engine 6 during flight.

[0032] In one embodiment, the first test chamber 2 includes an outer shell 21 and an exhaust duct 22, the outer shell 21 circumferentially surrounds a receiving chamber 20, the exhaust duct 22 and the propeller 5 are respectively located at two ends of the outer shell 21, and the exhaust duct 22 is connected to the outer shell 21, so that the flow channel in the exhaust duct 22 is connected to the receiving chamber 20. It should be noted that the outer shape structure of the exhaust duct 22 is a trapezoidal structure, and other structures besides the trapezoidal structure can also be other structures, such as rectangular structures. This facilitates the adaptation of the first test chamber 2 to meet the installation of turboshaft engines 6 with different intake and exhaust flow channel sizes.

[0033] In one embodiment, the mounting portion 7 includes a main fulcrum mounting frame 71 and a secondary fulcrum mounting frame 72, and the engine 6 is mounted on the main fulcrum mounting frame 71 and the secondary fulcrum mounting frame 72, and the main fulcrum mounting frame 71 and the secondary fulcrum mounting frame 72 are arranged on the cavity wall of the accommodating cavity 20 at a predetermined interval. It should be noted that the main fulcrum mounting frame 71 is at the center position of the engine 6, and the secondary fulcrum mounting frame 72 is at the edge position of the engine 6, and the main fulcrum mounting frame 71 and the secondary fulcrum mounting frame 72 are used to realize the installation structure layout of the simulated turboshaft engine 6 on the helicopter.

[0034] In one embodiment, a distortion plate is provided at the connection between the exhaust duct 22 and the engine 6. The distortion plate is further used to simulate the inlet airflow pressure distortion of the turboshaft engine 6. Specifically, an air inlet hole is provided on a part of the distortion plate, so as to implement the experiment of the inlet airflow pressure distortion. It should be noted that when the distortion plate is provided, the distortion test part (24) needs to be removed.

[0035] In one embodiment, a power interface and a gas circuit interface are provided on the first test cabin 2, and the power interface or the gas circuit interface cooperates with a corresponding structural interface on the first power pack 3. The power interface, the gas circuit interface and the transmission structure interface are matched to realize the same installation layout of the simulated turboshaft engine 6 on the helicopter. In addition, the gas source or power supply for starting the turboshaft engine 6 can be extracted from an independent power pack, and the transmission is realized through the power interface and the gas circuit interface to meet the starting methods of different turboshaft engines 6.

[0036] In one embodiment, a live-fire shooting device is provided outside the first test cabin 2. Optionally, the live-fire shooting device can also be fixed at other mounting positions of the aircraft 1, and the high-temperature exhaust gas of the ammunition fired by the live-fire shooting device can be absorbed by the air inlet of the power test cabin to simulate the temperature distortion of the inlet airflow of the turboshaft engine 6.

[0037] In one embodiment, a second test cabin 8 and a second power pack 9 are further included. The second test cabin 8 and the second power pack 9 are located on the same side of the aircraft 1, and the second test cabin 8 is symmetrical to the first test cabin 2 about the centerline of the aircraft 1, and the second power pack 9 is symmetrical to the first power pack 3 about the centerline of the aircraft 1. Optionally, the first test cabin 2 and the second test cabin 8 have the same structure, and the first power pack 3 and the second power pack 9 have the same structure. Optionally, the two power packs can be installed in the tail layout or in the two sides of the wing layout, and the starting method of the turboshaft engine 6 can take into account both the motor starting and the air turbine starter modes. Since the two test cabins and the two power packs have the same structure, the functions and effects are also the same, and the role and effect of the second test cabin 8 and the second power pack 9 will not be described here.

[0038] like Figure 3 As shown, it is a control system of a general test flight platform for a turboshaft engine 6, which controls the pitch control signal of the engine 6 in the test chamber, thereby controlling the pitch of the propeller 5, and then realizing the speed of the device, and realizing the speed matching control of the propeller 5 by the independent reduction box 4 by controlling the output power of the engine 6. The power pack is used to provide a starting air source or a starting current to the engine 6 in the test chamber. By controlling the ammunition firing of the live-fire shooting equipment, the heat generated enters the air inlet to realize temperature distortion. The simulated temperature distortion can also be realized by installing a distortion plate on the air inlet 23 of the engine 6. The distortion test section is controlled to realize the injection of water flow in the air inlet, and the temperature of the water flow is based on the environmental icing test of the simulated turboshaft engine 6, etc.

[0039] In summary, the universal flight test platform of the present invention is more likely to meet the flight envelope requirements of the turboshaft engine 6 design than the helicopter flight platform. In addition, the two test cabins can simulate the installation structure of the turboshaft engine 6 in the prototype helicopter without changing the control method of the turboshaft engine 6, and are more commonly used to carry out various types of turboshaft engine flight tests.

[0040] Although the present invention has been described in detail with reference to the aforementioned embodiments, those skilled in the art should understand that they can still modify the technical solutions described in the aforementioned embodiments, or make equivalent substitutions for some of the technical features therein; and these modifications or substitutions do not make the essence of the corresponding technical solutions deviate from the spirit and scope of the technical solutions of the embodiments of the present invention.

Claims

1. A universal flight test platform for a turboshaft engine, characterized in that: include: An aircraft (1), a first test cabin (2), a first power pack (3), an independent reduction box (4) and a propeller (5), wherein two wings are respectively arranged on both sides of the aircraft (1), the first test cabin (2) is arranged on one of the wings, the first test cabin (2) and the first power pack (3) are located on the same side of the aircraft (1), the first power pack (3) is arranged at the connection between one of the wings and the aircraft (1), the first test cabin (2) has a receiving cavity (20), and a The mounting portion (7) is used to mount the engine (6) and make the axial direction of the engine (6) consistent with that of the accommodating chamber (20); the engine (6) is connected to the first power pack (3) through the first test chamber (2); the independent reduction gearbox (4) is arranged in the accommodating chamber (20); the propeller (5) is arranged outside the accommodating chamber (20) and is located at the end of the first test chamber (2); the engine (6) is sequentially connected to the independent reduction gearbox (4) and the propeller (5) by transmission.

2. The universal flight test platform for turboshaft engines according to claim 1, characterized in that: An air intake passage (23) is provided in the accommodating chamber (20), one end of the air intake passage (23) passes through the accommodating chamber (20) and communicates with the outside, and the other end communicates with a flow passage of the engine (6), and a portion of the air intake passage (23) is between the engine (6) and the independent reduction box (4).

3. The universal flight test platform for turboshaft engines according to claim 2, characterized in that: A distortion test section (24) is provided between the intake passage (23) and a flow passage of the engine (6).

4. The universal flight test platform for turboshaft engines according to claim 2, characterized in that: A water spraying and freezing test section (25) is also provided in the accommodating chamber (20), and the water spraying and freezing test section (25) is connected to the air inlet (23) via a pipeline.

5. The universal flight test platform for turboshaft engines according to claim 1, characterized in that: The first test chamber (2) comprises an outer shell (21) and an exhaust duct (22); the outer shell (21) circumferentially surrounds the accommodating chamber (20); the exhaust duct (22) and the propeller (5) are respectively located at two ends of the outer shell (21); and the exhaust duct (22) is connected to the outer shell (21) so that a flow channel in the exhaust duct (22) is in communication with the accommodating chamber (20).

6. The universal flight test platform for turboshaft engines according to claim 1, characterized in that: The mounting portion (7) comprises a main fulcrum mounting frame (71) and a secondary fulcrum mounting frame (72), the engine being mounted on the main fulcrum mounting frame (71) and the secondary fulcrum mounting frame (72), and the main fulcrum mounting frame (71) and the secondary fulcrum mounting frame (72) being arranged on the cavity wall of the accommodating cavity (20) at a predetermined interval.

7. The universal flight test platform for turboshaft engines according to claim 5, characterized in that: A distortion plate is provided at the connection between the air inlet (23) and the engine (6).

8. The universal flight test platform for turboshaft engines according to claim 1, characterized in that: The first test chamber (2) is provided with an electric power interface and an air circuit interface, and the electric power interface or the air circuit interface cooperates with a corresponding structural interface on the first power pack (3).

9. The universal flight test platform for turboshaft engines according to claim 1, characterized in that: A live-fire shooting device is provided outside the first test chamber (2).

10. The universal flight test platform for a turboshaft engine according to any one of claims 1 to 9, characterized in that: It also includes a second test chamber (8) and a second power pack (9), wherein the second test chamber (8) and the second power pack (9) are located on the same side of the aircraft (1), and the second test chamber (8) is symmetrical to the first test chamber (2) with respect to the center line of the aircraft (1), and the second power pack (9) is symmetrical to the first power pack (3) with respect to the center line of the aircraft (1).

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