阅读说明：本技术 用于高超声速飞行器的大流量反向喷流试验装置及其方法 (High-flow reverse jet test device and method for hypersonic aircraft ) 是由 邱华诚 杨彦广 石义雷 李�杰 龙正义 孙良宝 贺江峰 强慢 于 2021-09-30 设计创作，主要内容包括：本发明公开了一种用于高超声速飞行器的大流量反向喷流试验装置及其方法。该试验装置包括若干个均匀分布在轨控舱的迎风面上、内部设置有拉瓦尔喷管、迎向来流的试验喷嘴；包括用于支撑模型和天平,并提供进气通道的天平支杆；还包括天平以及用于保护天平的隔热套Ⅰ、隔热套Ⅱ。高压空气从进气口进入气流通道,再分别通过支路管道从对应的试验喷嘴喷出。该试验方法通过专用的试验模型和天平测量气动力,拍摄流场显示照片,全面评估大流量反向喷流的技术效果。(The invention discloses a large-flow reverse jet test device and method for a hypersonic aircraft. The test device comprises a plurality of test nozzles which are uniformly distributed on the windward side of the rail control cabin, and inside which Laval spray pipes and oncoming inflow are arranged; comprises a balance support rod for supporting the model and the balance and providing an air inlet channel; the balance further comprises a balance, and a heat insulation sleeve I and a heat insulation sleeve II which are used for protecting the balance. High-pressure air enters the air flow channel from the air inlet and is sprayed out from the corresponding test nozzles through the branch pipelines respectively. The test method measures aerodynamic force through a special test model and a balance, takes a flow field display picture, and comprehensively evaluates the technical effect of large-flow reverse jet flow.)

1. A large-traffic reverse jet test device for hypersonic aircraft, characterized in that, large-traffic reverse jet test device:

the device comprises a plurality of test nozzles (6), wherein the test nozzles (6) are arranged between a model head (1) and a rail control cabin (11) and are uniformly distributed on the windward side of the rail control cabin (11) along the circumferential direction, the test nozzles (6) face to incoming flow, a Laval spray pipe located on the central axis of the test nozzles (6) is arranged at the front end of each test nozzle (6), and the rear ends of the test nozzles (6) are arranged in corresponding through holes in the windward side of the rail control cabin (11) in a cylindrical surface matching mode;

comprises a balance (4), wherein the balance (4) is a rod type six-component balance;

comprises a balance supporting rod (3), the balance supporting rod (3) is divided into a front section and a rear section, the front section and the rear section are isolated by a block, the front end of the balance supporting rod (3) is fixedly connected with a tail cone of a balance (4) by a positioning key (8), the rear end of the balance supporting rod (3) is fixed on a middle support of the high supersonic speed wind tunnel, the rear section of the balance supporting rod (3) is provided with an airflow channel on the central axis of the balance supporting rod (3), the position of the rear section of the balance supporting rod (3) close to the middle support is provided with an air inlet (2), the position of the rear section of the balance supporting rod (3) close to the block is provided with branch pipelines which correspond to test nozzles (6) one by one, and the branch pipelines are positioned in an inner cavity of a rail control cabin (11), the branch pipelines extend into the rear section of the balance support rod (3), high-pressure air enters the airflow channel from the air inlet (2) and is respectively sprayed out from the corresponding test nozzles (6) through the branch pipelines;

the heat insulation device comprises a heat insulation sleeve I (5), wherein the heat insulation sleeve I (5) is a taper sleeve and is sleeved on the front conical surface of a balance (4), and a model head (1) is sleeved on the front conical surface of the balance (4);

the balance with the heat insulation sleeve is characterized by further comprising a heat insulation sleeve II (7), the heat insulation sleeve II (7) is a stepped cylinder matched with the balance (4) measuring section and the tail cone in shape, the stepped cylinder is sleeved on the balance (4) measuring section and the tail cone, and the heat insulation sleeve II (7) is connected with the front end of the balance support rod (3) in a sealing mode.

2. The mass flow reverse jet test device for the hypersonic aircraft according to claim 1, characterized in that the included angle α between the central axis of the test nozzle (6) and the incoming flow is 25 ° -35 °.

3. The mass-flow reverse jet test device for the hypersonic aircraft according to claim 1, characterized in that more than one of the test nozzles (6) is replaced by a debugging nozzle (9), the front end of the debugging nozzle (9) is a boss, the rear end of the debugging nozzle is the same as the rear end of the test nozzle (6), a laval nozzle pipe located on the central axis of the debugging nozzle (9) is arranged in the front end of the boss of the debugging nozzle (9), and a pressure measuring hole (10) for measuring total pressure of branch pipeline jet flow is arranged at the rear end of the boss of the debugging nozzle (9).

4. The high-flow reverse jet flow test device for the hypersonic aircraft according to claim 1, characterized in that the test nozzle (6) is provided with a special four-claw tool, the front end face of the test nozzle (6) is provided with 4 counter bores which are centrosymmetric, the front end of the four-claw tool is provided with 4 cylindrical heads which are in one-to-one correspondence with the 4 counter bores, the 4 cylindrical heads respectively extend into the 4 counter bores, the four-claw tool grasps the test nozzle (6), and the test nozzle (6) is installed or pulled out of the branch pipeline.

5. The test method of the large flow reverse jet flow for the hypersonic aerocraft is based on the test device of the large flow reverse jet flow for the hypersonic aerocraft in claim 4, and is characterized by comprising the following steps:

a. determining the shrinkage ratio of a test model according to the blockage requirement of the hypersonic wind tunnel, designing and processing a shrinkage ratio jet flow test model of the hypersonic aircraft, and carrying out equal-proportion simulation on a model head (1) and an orbit control cabin (11);

b. designing and processing a balance (4) arranged in the shrinkage ratio jet flow test model, wherein the balance (4) is a rod type six-component balance, and the measurement precision of the axial force is 0.3%;

c. designing a four-claw tool for machining the test nozzle (6);

d. designing and processing a plurality of groups of test nozzles (6) and corresponding debugging nozzles (9), wherein the Mach numbers of the Laval nozzles in each group of test nozzles (6) and debugging nozzles (9) are the same;

e. installing a test model in the hypersonic wind tunnel, installing a flow field display device outside the hypersonic wind tunnel, and carrying out ground debugging;

f. developing a hypersonic wind tunnel test;

f1. the method comprises the following steps of performing ground test, wherein a first group of test nozzles (6) and corresponding debugging nozzles (9) are arranged on a model head (1), starting an external high-pressure air source, measuring total pressure of jet flow by pressure measuring holes (10) of the debugging nozzles (9), achieving a preset series of jet flow pressures, and recording corresponding series of regulating valve opening degrees;

f2. replacing the first set of test nozzles (6) with the first set of test nozzles (9);

f3. starting the hypersonic wind tunnel;

f4. after the wind tunnel inflow is stable, the high-pressure air source is opened according to the opening degree of the regulating valve, and the test nozzle (6) starts to jet;

f5. a balance (4) measures force, and a flow field display device shoots;

f6. stopping the hypersonic wind tunnel;

f7. evaluating the contribution of the reverse jet flow to shortening the test course of the flight test of the hypersonic aircraft through force measurement, pressure measurement and flow field display data, changing the opening degree of a regulating valve, adjusting the total pressure of jet flow, repeating the steps f 3-f 6 to carry out the hypersonic wind tunnel test, and completing the test of the first group of test nozzles (6);

f8. and (3) replacing another group of test nozzles (6) with different Mach numbers, and repeating the steps f 1-f 6 until the test of the last group of test nozzles (6) is completed.

6. The mass flow reverse jet test method for the hypersonic aircraft according to claim 5, characterized in that the flow field display device is a shadow system, a schlieren system or a glow display system.

Technical Field

The invention belongs to the technical field of hypersonic wind tunnel tests, and particularly relates to a high-flow reverse jet flow test device and method for a hypersonic aircraft.

Background

When the flight test of the hypersonic aircraft is carried out, if the required test flight distance is far, the test flight in the national border in the whole process cannot be realized. At present, it is urgently needed to develop a large-flow reverse jet test device and a large-flow reverse jet test method for a hypersonic aircraft, wherein the test aircraft is provided with a reverse jet device, and when flight test data of the hypersonic aircraft are obtained, the large-flow reverse jet test device is immediately started to decelerate the hypersonic aircraft so as to shorten a test voyage.

Disclosure of Invention

The invention aims to provide a high-flow reverse jet test device for a hypersonic aircraft, and the invention also aims to provide a high-flow reverse jet test method for the hypersonic aircraft.

The invention relates to a large-flow reverse jet test device for a hypersonic aircraft, which is characterized in that:

the test nozzle is arranged between the head of the model and the rail control cabin and is uniformly distributed on the windward side of the rail control cabin along the circumferential direction, the test nozzle faces the incoming flow, a Laval nozzle positioned on the central axis of the test nozzle is arranged at the front end of the test nozzle, and the rear end of the test nozzle is arranged in a corresponding through hole on the windward side of the rail control cabin in a cylindrical surface matching mode;

comprises a balance, wherein the balance is a rod type six-component balance;

the device comprises a balance supporting rod, wherein the balance supporting rod is divided into a front section and a rear section, the front section and the rear section are isolated by a blocking block, the front end of the balance supporting rod is fixedly connected with a balance tail cone through a positioning key, the rear end of the balance supporting rod is fixed on a middle support of the hypersonic wind tunnel, the rear section of the balance supporting rod is provided with an airflow channel on the central axis of the balance supporting rod, the rear section of the balance supporting rod is provided with an air inlet at a position close to the middle support, branch pipelines which are in one-to-one correspondence with test nozzles are arranged at a position close to the blocking block at the rear section of the balance supporting rod, the branch pipelines are positioned in an inner cavity of a rail control cabin, the branch pipelines extend into the rear section of the balance supporting rod, high-pressure air enters the airflow channel from the air inlet and is respectively sprayed out from the corresponding test nozzles through the branch pipelines;

the heat insulation device comprises a heat insulation sleeve I, wherein the heat insulation sleeve I is a taper sleeve and is sleeved on the front conical surface of the balance, and the head of a model is sleeved on the front conical surface of the balance;

the balance comprises a balance support rod and a heat insulation sleeve II, wherein the heat insulation sleeve II is a stepped cylinder matched with the balance measuring section and the tail cone in shape, the stepped cylinder is sleeved on the balance measuring section and the tail cone, and the heat insulation sleeve II is connected with the front end of the balance support rod in a sealing mode.

Further, the included angle alpha between the central axis of the test nozzle and the incoming flow ranges from 25 degrees to 35 degrees.

Furthermore, any one or more than one of the test nozzles are replaced by a debugging nozzle, the front end of the debugging nozzle is a boss, the rear end of the debugging nozzle is the same as the rear end of the test nozzle, a Laval nozzle positioned on the central axis of the debugging nozzle is arranged in the front end of the boss of the debugging nozzle, and a pressure measuring hole for measuring the total pressure of the branch pipeline jet flow is arranged at the rear end of the boss of the debugging nozzle.

Furthermore, the test nozzle be provided with dedicated four-claw frock, the preceding terminal surface of test nozzle is opened has 4 counter bores of centrosymmetry, the front end of four-claw frock is provided with 4 cylinder heads with 4 counter bores one-to-one, 4 cylinder heads stretch into 4 counter bores respectively, four-claw frock grasps test nozzle, with test nozzle installation or extract branch road pipeline.

The invention discloses a large-flow reverse jet test method for a hypersonic aircraft, which comprises the following steps of:

a. determining the shrinkage ratio of a test model according to the blockage degree requirement of the hypersonic wind tunnel, designing and processing a shrinkage ratio jet flow test model of the hypersonic aircraft, and carrying out equal-proportion simulation on the head of the model and the rail control cabin;

b. designing and processing a balance arranged in a shrinkage ratio jet flow test model, wherein the balance is a rod type six-component balance, and the measurement precision of the axial force is 0.3%;

c. designing a four-claw tool for machining a test nozzle;

d. designing and processing a plurality of groups of test nozzles and corresponding debugging nozzles, wherein the Mach numbers of the Laval nozzles in each group of test nozzles and debugging nozzles are the same;

e. installing a test model in the hypersonic wind tunnel, installing a flow field display device outside the hypersonic wind tunnel, and carrying out ground debugging;

f. developing a hypersonic wind tunnel test;

f1. the method comprises the following steps of performing ground test, namely installing a first group of test nozzles and corresponding debugging nozzles at the head of a model, starting an external high-pressure air source, measuring total jet pressure by pressure measuring holes of the debugging nozzles to achieve a preset series of jet pressure, and recording the corresponding series of opening degrees of regulating valves;

f2. replacing the first group of debugging nozzles with testing nozzles;

f3. starting the hypersonic wind tunnel;

f4. after the wind tunnel inflow is stable, opening a high-pressure air source according to the opening degree of the regulating valve, and starting jet flow by the test nozzle;

f5. balance force measurement and flow field display device shooting;

f6. stopping the hypersonic wind tunnel;

f7. evaluating the contribution of the reverse jet flow to shortening the test course of the flight test of the hypersonic aircraft through force measurement, pressure measurement and flow field display data, changing the opening degree of a regulating valve, adjusting the total pressure of jet flow, repeating the steps f 3-f 6 to carry out the hypersonic wind tunnel test, and completing the test of the first group of test nozzles;

f8. and (4) replacing another group of test nozzles with different Mach numbers, and repeating the steps f 1-f 6 until the test of the last group of test nozzles is completed.

Further, the flow field display device is one of a shadow system, a schlieren system or a glow display system.

The high-flow reverse jet test device for the hypersonic aircraft adopts the test nozzles uniformly distributed on the windward side of the rail control cabin along the circumferential direction to carry out high-flow reverse jet, can carry out pneumatic braking to the maximum extent, and comprehensively evaluates the feasibility of the pneumatic braking through force measurement, pressure measurement and flow field display data.

The high-flow reverse jet flow test device for the hypersonic aircraft is provided with the special four-claw tool for mounting and dismounting the test nozzle and the debugging nozzle, and aims to realize the quick mounting and dismounting of the test nozzle and the debugging nozzle in a narrow space and improve the test efficiency.

The balance used in the high-flow reverse jet test device of the hypersonic aircraft has high axial force measurement precision, and can accurately evaluate the influence of 'pneumatic braking' on the axial force.

The balance used in the high-flow reverse jet test device of the hypersonic aircraft has good protection measures, the balance and the high-heat incoming flow of the hypersonic wind tunnel are isolated and protected through the heat insulation sleeve I and the heat insulation sleeve II, and adverse effects of the temperature effect of the balance on measurement are avoided as much as possible.

The branch pipeline used in the high-flow reverse jet test device of the hypersonic aircraft has the function of a back-jet engine chamber, the volume of the back-jet engine chamber is adjusted by adjusting the inner diameter of the branch pipeline, the flow speed of high-pressure air in the branch pipeline is reduced as much as possible, and the high-pressure air can be made to be uniform and stable as much as possible before entering the test nozzle.

The high-flow reverse jet test method for the hypersonic aircraft, disclosed by the invention, is used for simulating the appearance of the hypersonic aircraft, simulating the incoming flow and simulating the jet, can be used for comprehensively evaluating the high-flow reverse jet of the hypersonic aircraft, obtains comprehensive force measurement, pressure measurement and flow field display data, and has the advantages of fine simulation and accurate evaluation.

Drawings

FIG. 1 is a schematic perspective view of a high flow reverse jet test apparatus for a hypersonic aircraft according to the present invention;

FIG. 2 is a partial cross-sectional view of a balance and a balance strut of the mass flow reverse jet test apparatus for a hypersonic aircraft according to the present invention;

FIG. 3 is a schematic view of the installation of a test nozzle and a debugging nozzle in the large-flow reverse jet test device for the hypersonic aircraft according to the invention;

FIG. 4 is an included angle α between the central axis of the test nozzle and the incoming flow in the large-flow reverse jet test device for the hypersonic aircraft according to the present invention;

FIG. 5 is a schematic perspective view of a pilot nozzle of the present invention for use in a high flow reverse jet test apparatus for hypersonic aircraft;

fig. 6 is a perspective sectional view of a debugging nozzle in the large flow reverse jet test device for the hypersonic aircraft according to the present invention.

In the figure, 1. model head; 2. an air inlet; 3. a balance support bar; 4. a balance; 5. a heat insulation sleeve I; 6. a test nozzle; 7. a heat insulation sleeve II; 8. a positioning key; 9. debugging a nozzle; 10. a pressure measuring hole; 11. and a rail control cabin.

Detailed Description

The present invention will be described in detail below with reference to the accompanying drawings and examples.

As shown in fig. 1, 2, and 3, the high flow rate reverse jet test apparatus for a hypersonic aircraft according to the present invention is characterized in that:

the test device comprises a plurality of test nozzles 6, wherein the test nozzles 6 are arranged between a model head 1 and a rail control cabin 11 and are uniformly distributed on the windward side of the rail control cabin 11 along the circumferential direction, the test nozzles 6 face the incoming flow, a Laval nozzle located on the central axis of the test nozzles 6 is arranged at the front ends of the test nozzles 6, and the rear ends of the test nozzles 6 are arranged in corresponding through holes on the windward side of the rail control cabin 11 in a cylindrical surface matching mode;

comprises a balance 4, wherein the balance 4 is a rod type six-component balance;

the high-pressure air flow test device comprises a balance supporting rod 3, wherein the balance supporting rod 3 is divided into a front section and a rear section, the front section and the rear section are isolated by a blocking block, the front end of the balance supporting rod 3 is fixedly connected with a tail cone of a balance 4 through a positioning key 8, the rear end of the balance supporting rod 3 is fixed on a middle support of a hypersonic wind tunnel, the rear section of the balance supporting rod 3 is provided with an air flow channel on the central axis of the balance supporting rod 3, the rear section of the balance supporting rod 3 close to the middle support is provided with an air inlet 2, the rear section of the balance supporting rod 3 close to the blocking block is provided with branch pipelines which correspond to test nozzles 6 one by one, the branch pipelines are positioned in an internal cavity of a rail control cabin 11, the branch pipelines extend into the rear section of the balance supporting rod 3, high-pressure air enters the air flow channel from the air inlet 2 and is respectively sprayed out from the corresponding test nozzles 6 through the branch pipelines;

the heat insulation device comprises a heat insulation sleeve I5, wherein the heat insulation sleeve I5 is a taper sleeve and is sleeved on the front conical surface of a balance 4, and a model head 1 is sleeved on the front conical surface of the balance 4;

the balance comprises a balance 4, and is characterized by further comprising a heat insulation sleeve II 7, wherein the heat insulation sleeve II 7 is a stepped cylinder matched with the measuring section of the balance 4 and the outer shape of a tail cone, the stepped cylinder is sleeved on the measuring section of the balance 4 and the tail cone, and the heat insulation sleeve II 7 is connected with the front end of the balance support rod 3 in a sealing mode.

Further, as shown in fig. 4, the included angle α between the central axis of the test nozzle 6 and the incoming flow is in the range of 25 ° to 35 °.

Further, as shown in fig. 5 and 6, any one or more than one of the test nozzles 6 is replaced by a debugging nozzle 9, the front end of the debugging nozzle 9 is a boss, the rear end of the debugging nozzle is the same as the rear end of the test nozzle 6, a laval nozzle located on the central axis of the debugging nozzle 9 is arranged in the front end of the boss of the debugging nozzle 9, and a pressure measuring hole 10 for measuring total pressure of branch pipeline jet flow is arranged at the rear end of the boss of the debugging nozzle 9.

Furthermore, the test nozzle 6 is provided with a special four-claw tool, 4 counter bores with central symmetry are formed in the front end face of the test nozzle 6, 4 cylindrical heads in one-to-one correspondence with the 4 counter bores are arranged at the front end of the four-claw tool, the 4 cylindrical heads respectively extend into the 4 counter bores, the four-claw tool tightly grasps the test nozzle 6, and the test nozzle 6 is installed or pulled out of the branch pipeline.

The invention discloses a large-flow reverse jet test method for a hypersonic aircraft, which comprises the following steps of:

a. determining the shrinkage ratio of a test model according to the blockage degree requirement of the hypersonic wind tunnel, designing and processing a shrinkage ratio jet flow test model of the hypersonic aircraft, and carrying out equal-proportion simulation on a model head 1 and an orbit control cabin 11;

b. designing and processing a balance 4 arranged in the shrinkage ratio jet flow test model, wherein the balance 4 is a rod type six-component balance, and the measurement precision of the axial force is 0.3%;

c. designing a four-claw tool for machining the test nozzle 6;

d. designing and processing a plurality of groups of test nozzles 6 and corresponding debugging nozzles 9, wherein the Mach numbers of the Laval nozzles in each group of test nozzles 6 and debugging nozzles 9 are the same;

e. installing a test model in the hypersonic wind tunnel, installing a flow field display device outside the hypersonic wind tunnel, and carrying out ground debugging;

f. developing a hypersonic wind tunnel test;

f1. the ground test comprises the steps that a first group of test nozzles 6 and corresponding debugging nozzles 9 are installed on a model head 1, an external high-pressure air source is started, pressure measuring holes 10 of the debugging nozzles 9 measure total pressure of jet flow to reach a preset series of jet flow pressures, and corresponding series of opening degrees of adjusting valves are recorded;

f2. replacing the first set of commissioning nozzles 9 with testing nozzles 6;

f3. starting the hypersonic wind tunnel;

f4. after the wind tunnel inflow is stable, the high-pressure air source is opened according to the opening degree of the regulating valve, and the test nozzle 6 starts to jet;

f5. a balance 4 measures force, and a flow field display device shoots;

f6. stopping the hypersonic wind tunnel;

f7. evaluating the contribution of the reverse jet flow to shortening the test course of the flight test of the hypersonic aircraft through force measurement, pressure measurement and flow field display data, changing the opening degree of a regulating valve, adjusting the total pressure of jet flow, repeating the steps f 3-f 6 to carry out the hypersonic wind tunnel test, and completing the test of the first group of test nozzles 6;

f8. and (4) replacing another group of test nozzles 6 with different Mach numbers, and repeating the steps f 1-f 6 until the test of the last group of test nozzles 6 is completed.

Further, the flow field display device is one of a shadow system, a schlieren system or a glow display system.

Example 1

The large-flow reverse jet test device for the hypersonic aircraft in the embodiment is provided with 4 test nozzles 6 and 1 debugging nozzle 9, and any one test nozzle 6 can be replaced by the debugging nozzle 9.

According to the large-flow reverse jet flow test method for the hypersonic aircraft, high-pressure air is used as a jet flow medium, the wind tunnel wall air source adapter is externally connected with a high-pressure air source, the wind tunnel wall air source adapter, a phi 15mm interface on the middle support and the air inlet 2 are connected through a phi 19 multiplied by 2mm copper tube, and the high-pressure air enters the air flow channel from the air inlet 2 and is respectively sprayed out from the corresponding test nozzles 6 through branch pipelines. The ventilation area of the back-spray engine chamber formed by each branch pipeline is 78.5mm²The throat area of the Laval nozzle of the test nozzle 6 is 15.9mm²The flow velocity of the high-pressure air in the branch duct is about 40 m/s.

The embodiment adjusts the flow rate of the pipeline by changing the pressure of the high-pressure air source.

Although the embodiments of the present invention have been disclosed, the embodiments are not limited to the applications listed in the description and the embodiments, and can be fully applied to various fields of hypersonic boundary layer transition mode methods suitable for the present invention. Additional modifications and refinements of the present invention will readily occur to those skilled in the art without departing from the principles of the present invention, and therefore the present invention is not limited to the specific details and illustrations shown and described herein without departing from the general concept defined by the claims and their equivalents.