CN115436010B - Jet pipe thrust measurement test method based on afterbody and jet pipe integrated design - Google Patents

Jet pipe thrust measurement test method based on afterbody and jet pipe integrated design Download PDF

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CN115436010B
CN115436010B CN202211388072.9A CN202211388072A CN115436010B CN 115436010 B CN115436010 B CN 115436010B CN 202211388072 A CN202211388072 A CN 202211388072A CN 115436010 B CN115436010 B CN 115436010B
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thrust
outflow
effective
spray pipe
nozzle
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CN115436010A (en
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张诣
曾利权
李耀华
熊能
李建强
苗磊
尹疆
梁锦敏
苏博
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High Speed Aerodynamics Research Institute of China Aerodynamics Research and Development Center
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High Speed Aerodynamics Research Institute of China Aerodynamics Research and Development Center
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    • G—PHYSICS
    • G01—MEASURING; TESTING
    • G01M—TESTING STATIC OR DYNAMIC BALANCE OF MACHINES OR STRUCTURES; TESTING OF STRUCTURES OR APPARATUS, NOT OTHERWISE PROVIDED FOR
    • G01M9/00—Aerodynamic testing; Arrangements in or on wind tunnels
    • G01M9/06—Measuring arrangements specially adapted for aerodynamic testing
    • G01M9/062—Wind tunnel balances; Holding devices combined with measuring arrangements
    • G—PHYSICS
    • G01—MEASURING; TESTING
    • G01M—TESTING STATIC OR DYNAMIC BALANCE OF MACHINES OR STRUCTURES; TESTING OF STRUCTURES OR APPARATUS, NOT OTHERWISE PROVIDED FOR
    • G01M9/00—Aerodynamic testing; Arrangements in or on wind tunnels
    • G01M9/02—Wind tunnels
    • G01M9/04—Details
    • G—PHYSICS
    • G01—MEASURING; TESTING
    • G01M—TESTING STATIC OR DYNAMIC BALANCE OF MACHINES OR STRUCTURES; TESTING OF STRUCTURES OR APPARATUS, NOT OTHERWISE PROVIDED FOR
    • G01M9/00—Aerodynamic testing; Arrangements in or on wind tunnels
    • G01M9/08—Aerodynamic models
    • Y—GENERAL TAGGING OF NEW TECHNOLOGICAL DEVELOPMENTS; GENERAL TAGGING OF CROSS-SECTIONAL TECHNOLOGIES SPANNING OVER SEVERAL SECTIONS OF THE IPC; TECHNICAL SUBJECTS COVERED BY FORMER USPC CROSS-REFERENCE ART COLLECTIONS [XRACs] AND DIGESTS
    • Y02—TECHNOLOGIES OR APPLICATIONS FOR MITIGATION OR ADAPTATION AGAINST CLIMATE CHANGE
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Abstract

The invention belongs to the technical field of thrust vector control of aircrafts, and discloses a jet pipe thrust measurement test method based on afterbody and jet pipe integrated design. The thrust measurement test method for the spray pipe comprises the following steps: designing and processing a wind tunnel test model; carrying out a tail support force measurement wind tunnel test; carrying out a measurement test that the head supporting spray pipe has thrust and outflow; carrying out a measurement test of thrust and outflow of the head support spray pipe; performing a non-thrust and outflow measurement test on the head support spray pipe; calculating the effective thrust characteristic of the test model in a jet flow state; calculating the disturbance aerodynamic characteristics of the outflow to the thrust characteristics; and calculating the aerodynamic characteristics of the full computer model with the jet flow state. The thrust-reducing-nozzle resistance data of the nozzle can be directly measured by the nozzle thrust measurement test method, authenticity and reliability of the data are improved, and the method is beneficial to reducing the risk of the integrated design technology of the flight and the research and development cost and shortening the development period.

Description

Nozzle thrust measurement test method based on afterbody and nozzle integrated design
Technical Field
The invention belongs to the technical field of thrust vector control of aircrafts, and particularly relates to a thrust measurement test method of a spray pipe based on the integrated design of a afterbody and the spray pipe.
Background
The thrust vector control technology is a technology for directly using a part of engine thrust for aircraft flight control through jet flow steering of a nozzle. The aircraft as a whole has the mutual influence and inseparability of the external circumfluence of the aircraft body and the internal fluence of the propulsion system. The inner flow is captured by the air inlet, and returns to the outer flow after being subjected to speed reduction and diffusion of the air inlet, combustion pressurization of the engine and accelerated pressure reduction of the spray pipe. This process generates a series of complex flow coupling interference phenomena, which cause coordination and matching problems among the air intake duct, the engine, the nozzle, and the Airframe, i.e., a Propulsion-Airframe Integration problem, and directly affect the aerodynamic, propulsion, maneuvering, and safety performance of the aircraft. The maximization of the internal and external flow pneumatic comprehensive benefits can be realized only by fully recognizing the flow coupling interference characteristics among the engine body/air inlet channel, the air inlet channel/engine and the spray pipe/engine body and clarifying the action rules and the influence characteristics of the engine body/air inlet channel/engine and the spray pipe/engine body, so that the compatibility of a propulsion system and the engine body is ensured, and the design target is reached. Among these, for the nozzle, in order to evaluate the thrust efficiency of the air jet engine exhaust under outflow conditions, it is necessary to measure the effective thrust of the nozzle, which is the most interesting indicator. The thrust characteristic of the nozzle is equal to the sum of the thrust of the nozzle (positive value) and the resistance acting on the nozzle shell (negative value), namely the thrust of the nozzle-the reduction-the resistance of the nozzle.
Usually, aircraft thrust vectoring nozzle characteristic data adopts the afterbody support or belly/back support mode, in order to record individual nozzle thrust, test model usually adopts two-layer nested method, separate interior nozzle and outside fuselage casing, must keep enough clearance between two-layer and prevent to collide with each other, lead to the outside streaming of organism and propulsion system internal flow to influence each other simulation distortion, and the internal cavity pressure between two-layer is difficult to survey accurately, the data of record is the nozzle thrust that does not deduct the resistance data, need can obtain the nozzle characteristic parameter of real concern of "nozzle thrust-subtract-nozzle resistance" through a series of revisions, thereby obtain the effective thrust characteristic under the accurate jet current state. And then, through a series of data processing, the disturbance aerodynamic characteristics of the outflow to the thrust characteristics and the aerodynamic characteristics of the full-engine model with the jet flow state can be further obtained.
At present, a thrust measurement test method of the spray pipe based on the integrated design of the afterbody and the spray pipe needs to be developed.
Disclosure of Invention
The invention aims to solve the technical problem of providing a test method for measuring thrust of a spray pipe based on the integrated design of a afterbody and the spray pipe.
The invention relates to a test method for measuring thrust of a spray pipe based on the integrated design of a afterbody and the spray pipe, which comprises the following steps of:
s10, designing and processing a wind tunnel test model;
the wind tunnel test model comprises a tail support force measurement test model and a head support spray pipe thrust measurement test model which are consistent in model scaling and rear body appearance; the measuring balance of the tail support force measuring test model is a rod balance; the measuring balance of the head support spray pipe thrust measuring test model is a ring balance, and the head support spray pipe thrust measuring test model is also provided with a spray pipe and a spray pipe block for replacing the spray pipe; the distance between a ring balance of the head support spray pipe thrust measurement test model and a rod balance calibration center of the tail support force measurement test model is as follows: the distance delta l in the X direction, the distance delta Y in the Y direction and the distance delta Z in the Z direction;
s20, carrying out a tail support force measurement wind tunnel test;
starting the wind tunnel, and measuring six-component aerodynamic force and aerodynamic moment F of the tail support force measurement test model through the rod balance under the conditions of set incoming flow Mach number and Reynolds number x 、F y 、F z 、m x 、m y 、m z ；
Wherein, F x Axial force for tail support, F y Normal force for tail support, F z Lateral force for tail support, m x Roll moment for tail support, m y Yaw moment, m, for tail support z A pitching moment for the tail support;
s30, performing a measurement test that the head support spray pipe has thrust and outflow;
installing a spray pipe on the head support spray pipe thrust measurement test model; opening the wind tunnel, under the conditions of incoming flow Mach number and Reynolds number which are the same as those in the step S20, conveying high-pressure airflow to the spray pipe by the external high-pressure air source, spraying the high-pressure airflow out of the spray pipe to generate thrust, and using the ring balance to balance the balanceMeasuring six-component aerodynamic force and aerodynamic moment F of afterbody under the condition of thrust and outflow x0 、F y0 、F z0 、m x0 、m y0 、m z0 ；
Wherein, F x0 For axial forces of the afterbody with thrust and outflow, F y0 Normal force of the rear body with thrust and outflow, F z0 For the lateral force of the rear body with thrust and outflow, m x0 Roll moment of the rear body under thrust and outflow conditions, m y0 For yawing moment m of the afterbody under the conditions of thrust and outflow z0 The pitching moment of the afterbody under the condition of thrust and outflow;
s40, performing a measurement test on whether the thrust of the head support spray pipe flows outwards or not;
reserving a nozzle on the head support nozzle thrust measurement test model; closing the wind tunnel, delivering the same high-pressure airflow to the spray pipe by an external high-pressure air source, spraying the high-pressure airflow out of the spray pipe to generate thrust, and measuring six-component aerodynamic force and aerodynamic moment F of the afterbody under the condition that the thrust does not flow outwards by the ring balance x01 、F y01 、F z01 、m x01 、m y01 、m z01 ；
Wherein, F x01 Axial force of the rear body with or without thrust and outflow, F y01 Normal force of the rear body with thrust and without outflow, F z01 M is the lateral force of the rear body under the condition of thrust and no outflow x01 Roll moment m of the rear body under the conditions of thrust and no outflow y01 The yawing moment m of the afterbody under the condition of thrust and no outflow z01 Pitching moment of the rear body under the condition of thrust and no outflow;
s50, performing a head support spray pipe non-thrust and outflow measurement test;
replacing the nozzle of the head support nozzle thrust measurement test model with a nozzle block; opening the wind tunnel, under the same incoming flow Mach number and Reynolds number conditions as the step S20, measuring six-component aerodynamic force and aerodynamic moment F of the afterbody under the conditions of existence of thrust and outflow by the ring balance x1 、F y1 、F z1 、m x1 、m y1 、m z1 ；
Wherein, F x1 Axial force of the rear body in the absence of thrust and with outflow, F y1 Normal force of the rear body in the absence of thrust and with outflow, F z1 The lateral force m of the rear body under the condition of no thrust and outflow x1 Roll moment m of the rear body in the absence of thrust and with outflow y1 Yaw moment m of the afterbody under the condition of no thrust and outflow z1 The pitching moment of the rear body under the condition of no thrust and outflow;
s60, calculating the effective thrust characteristic of the test model in a jet flow state;
calculating the effective thrust characteristic of the test model in the jet flow state according to the test data obtained in the steps S30 and S50:
effective aerodynamic force: f x is effective = F x0 - F x1 ，F y is effective = F y0 - F y1 ，F z is effective = F z0 - F z1 ；
Effective aerodynamic moment: m is x is effective =m x0 -m x1 ，m y is effective =m y0 -m y1 ，m z is effective =m z0 -m z1 ；
Wherein, F x is effective Effective axial thrust of the rear body, F y is effective Effective normal thrust of the rear body, F z is effective Effective side thrust of the rear body, m x is effective Effective roll moment of the rear body, m y is effective Effective yaw moment, m, of the rear body z is effective Effective pitching moment of the rear body;
s70, calculating the interference aerodynamic characteristics of the outflow to the thrust characteristics;
calculating the disturbance aerodynamic characteristics of the outflow to the thrust characteristics according to the test data obtained in the steps S30 and S40:
outflow disturbance aerodynamic force: f x interference = F x0 - F x01 ，F y interference = F y0 - F y01 ，F z interference = F z0 - F z01 ；
Outflow disturbance aerodynamic moment: m is x interference =m x0 -m x01 ，m y interference =m y0 -m y01 ，m z interference =m z0 -m z01 ；
Wherein, F x interference Axial force for disturbance of outflow to jet flow, F y interference Normal force for disturbance of outflow to jet flow, F z interference Lateral force for disturbance of outflow to jet flow, m x interference Roll moment m for disturbance of outflow to jet flow y interference Yawing moment m for disturbance of outflow on jet flow z interference The external flow is disturbed to the pitching moment of the jet flow;
s80, calculating the pneumatic characteristic of the full-computer model with the jet flow state;
and (5) correcting and correcting the tail support force measurement test data in the step (S10) by adopting the calculation result in the step (S60) to obtain the pneumatic characteristic of the jet flow state of the full-engine model:
jet state aerodynamic force: f x ’= F x + F x is effective ，F y ’= F y + F y is effective ，F z ’= F z + F z is effective ；
Jet state aerodynamic moment: m is x ’=m x +m x is effective + F y is effective ×Δz+ F z is effective ×Δy，m y ’=m y +m y is effective + F x is effective ×Δz+ F z is effective ×Δl，m z ’=m z +m z is effective + F x is effective ×Δy+ F y is effective ×Δl；
Wherein, F x ' axial force in full-machine-band jet state, F y ' Normal force with jet flow state for the whole machine, F z ' lateral force in full machine with jet flow state, m x ' Rolling moment m in the state of full machine with jet flow y ' yaw moment m in a full-machine band jet state z ' is the pitching moment of the whole machine with the jet flow state.
Furthermore, the tail support force measurement test model comprises a host model provided with a tail nozzle, a rod balance is arranged in an inner cavity of the host model, the rear end of the rod balance is connected with a tail support rod, and the rear end of the tail support rod is fixed on a middle support of the wind tunnel.
Furthermore, the head support spray pipe thrust measurement test model comprises a ventilating blade, a ventilating support rod, a ring balance, an aircraft afterbody, a spray pipe, a fairing and a corrugated pipe; the spray pipe plugging block is used for replacing the spray pipe;
the ventilation supporting rod is a circular tube body I, the front end of the ventilation supporting rod is closed, and the rear end of the ventilation supporting rod is opened;
a fairing is arranged on the end head of the front end of the ventilating strut; the circumferential direction of the front end of the ventilating support rod is provided with ventilating blades which are distributed in an axisymmetric way;
the rear end of the ventilating support rod is sequentially connected with the measuring section, the aircraft rear body and the spray pipe, the aircraft rear body is a circular pipe body II, and the spray pipe is an axisymmetric spray pipe; the measuring section comprises a ring type balance and a corrugated pipe which are sleeved in sequence from outside to inside, the fixed end of the ring type balance is installed on the outer layer installation surface at the rear end of the ventilating support rod, the free end of the ring type balance is installed on the outer layer installation surface at the front end of the rear body of the aircraft, the front end of the corrugated pipe is installed on the inner layer installation surface at the rear end of the ventilating support rod, and the rear end of the corrugated pipe is installed on the inner layer installation surface at the front end of the rear body of the aircraft;
the central axes of the fairing, the ventilating blade, the ventilating support rod, the ring balance, the aircraft rear body and the corrugated pipe are superposed with the central axis of the wind tunnel test section; an included angle alpha is formed between the central axis of the spray pipe and the central axis of the wind tunnel test section, and alpha is a spray pipe deflection angle;
a plurality of blade airflow channels communicated with an external high-pressure air source are arranged inside the ventilation blade, airflow outlets of the blade airflow channels are positioned on the inner wall of the ventilation support rod, and a ventilation support rod airflow channel is arranged on the central axis of the ventilation support rod; the high-pressure air source airflow enters the airflow channel of the ventilating support rod along the airflow channel of the blade through the airflow outlet and is sprayed out through the corrugated pipe and the spray pipe; the ring balance measures aerodynamic force and aerodynamic moment of the spray pipe; the corrugated pipe deforms along with the ring balance, and high-pressure air source airflow in the airflow channel of the ventilation support rod is sealed and isolated, so that the influence of the high-pressure air source airflow on the measurement of the ring balance is avoided.
Furthermore, the deflection angle alpha of the nozzle ranges from minus 20 degrees to 20 degrees.
Furthermore, the windward side of the ventilating vane is provided with an arc windward side for rectification.
Further, the front end of the ventilation supporting rod is closed by a blocking cover; the blocking cover is a step cylinder, and the diameter of the front section cylinder of the blocking cover is smaller than that of the rear section cylinder; the rear section of the cylinder of the blocking cover is inserted into the front end of the ventilation support rod; the front section cylinder of the blocking cover extends out of the ventilating support rod, and the clamping sleeve is sleeved on the front section cylinder of the blocking cover; the rear end of the cutting sleeve is provided with a circular groove matched with the front end head of the ventilation support rod, the rear end face of the cutting sleeve is a contact surface of the cutting sleeve and the front end face of the ventilation blade, and the front end of the cutting sleeve is provided with a locking nut; screwing the locking nut, tightly pushing the front end face of the ventilating blade by the clamping sleeve, and fixing the blocking cover and the clamping sleeve;
the fairing is sleeved on the clamping sleeve and wraps the blocking cover, the locking nut and the clamping sleeve, and the spherical outer surface of the fairing faces to incoming flow and is used for incoming flow rectification.
Furthermore, a sealing ring I is arranged on the inner wall of the cutting sleeve and used for sealing the contact surface of the cutting sleeve and the front section cylinder of the blocking cover; the rear end face of the clamping sleeve is provided with a sealing ring II for sealing the contact surface of the rear end face of the clamping sleeve and the front end face of the ventilating blade; a sealing ring III is arranged on the rear section of the cylinder of the blocking cover and used for sealing the contact surface of the blocking cover and the inner wall of the ventilating support rod; a sealing ring IV is arranged at the front section of the ventilation supporting rod and used for sealing the contact surface of the ventilation supporting rod and the rear end surface of the ventilation blade; the front end of the corrugated pipe is provided with a sealing ring V for sealing the contact surface between the inner layer mounting surface at the rear end of the ventilating support rod and the front end of the corrugated pipe, and the front end of the corrugated pipe is also provided with a sealing ring V for sealing the contact surface between the inner layer mounting surface at the front end of the rear body of the aircraft and the rear end of the corrugated pipe.
According to the head support spray pipe thrust measurement test model in the spray pipe thrust measurement test method based on the afterbody and spray pipe integrated design, an aircraft model obtained by scaling of an aircraft is simplified, the head of the aircraft model is removed, the rear section of the aircraft model and a ventilation support rod are designed into a whole, and finally, a circular pipe body of the ventilation support rod is adopted as the aircraft model. In the test process, parameters such as the flying speed, the height and the like of the aircraft are simulated by the incoming flow of the wind tunnel, and the incoming flow of the airflow is parallel to the central axis of the ventilating strut. Meanwhile, the deflection angle alpha of the spray pipe is changed to realize the steering jet flow of the spray pipe, and the spray pipe becomes a vector spray pipe; the high-pressure air source airflow enters the airflow channel of the ventilating support rod along the blade airflow channel, and is sprayed out through the corrugated pipe, the aircraft afterbody and the spray pipe to simulate the jet flow of the vectoring spray pipe. The corrugated pipe has the functions of ventilation but does not transmit force and moment, and does not influence the aerodynamic force and aerodynamic moment of the measuring nozzle of the ring balance.
According to the thrust measurement test method of the spray pipe based on the integrated design of the afterbody and the spray pipe, the afterbody support force measurement test model and the head support spray pipe thrust measurement test model are designed integrally with the spray pipe, so that the problems of simulation distortion caused by mutual influence of internal flow and external flow due to the separated design of the outer shell and the inner spray pipe of the test model, difficulty in measurement of pressure of an internal cavity and further difficulty in later-stage data processing caused by the difficulty in measurement are solved, the target data of thrust-reduction-spray pipe resistance of the spray pipe can be directly measured, the authenticity and reliability of data are improved, the later-stage data processing process is simplified, the risk and research and development cost of the integrated design of the aircraft engine are reduced, and the period of the aircraft engine is shortened.
Drawings
FIG. 1 is a flow chart of a nozzle thrust measurement test method based on a afterbody and nozzle integrated design of the present invention;
FIG. 2 is an installation schematic diagram of a tail support force measurement test model in the nozzle thrust measurement test method based on the afterbody and nozzle integrated design of the invention;
FIG. 3 is a schematic structural diagram (perspective view) of a head support nozzle thrust measurement test model in the nozzle thrust measurement test method based on the afterbody and nozzle integrated design of the present invention;
FIG. 4 is a schematic structural diagram (cross-sectional view) of a head support nozzle thrust measurement test model in the nozzle thrust measurement test method based on the afterbody and nozzle integrated design of the present invention;
FIG. 5 is a schematic structural diagram (a nozzle, a sectional view) of a vent strut head fairing of a head support nozzle thrust measurement test model in the nozzle thrust measurement test method based on the integrated design of the afterbody and the nozzle;
FIG. 6 is a schematic view (cross-sectional view) of a nozzle deflection angle of a head support nozzle thrust measurement test model in a nozzle thrust measurement test method based on a afterbody and nozzle integrated design according to the present invention;
FIG. 7 is a schematic structural view (nozzle block, cross-sectional view) of a vent strut head fairing of a head support nozzle thrust measurement test model in a nozzle thrust measurement test method based on an afterbody and nozzle integrated design of the present invention;
FIG. 8 is a schematic view (cross-sectional view) of the nozzle block installation of a head-supported nozzle thrust measurement test model in the nozzle thrust measurement test method based on the afterbody and nozzle integrated design of the present invention;
fig. 9 is a schematic structural diagram (cross-sectional view) of a nozzle blocking block of a head-supported nozzle thrust measurement test model in the nozzle thrust measurement test method based on the afterbody and nozzle integrated design of the present invention.
In the figure, 1, a gas collecting chamber; 2. a ventilation blade; 3. a ventilation strut; 4. a ring balance; 5. an aircraft aft body; 6. a nozzle; 7. a cowling; 8. blocking the cover; 9. locking the nut; 10. a card sleeve; 11. a sealing ring I; 12. a sealing ring II; 13. a sealing ring III; 14. a sealing ring IV; 15. a sealing ring V; 16. a vane airflow passage; 17. an airflow outlet; 18. a vent strut airflow channel; 19. a high pressure gas pipe inlet; 20. a bellows; 21. a nozzle block;
201. a host model; 202. a tail nozzle; 203. a bar balance; 204. a tail strut.
Detailed Description
The present invention will be described in detail below with reference to the accompanying drawings and examples.
Example 1
The nozzle thrust measurement test method based on the afterbody and nozzle integrated design is applied to the high-speed free jet wind tunnel.
As shown in fig. 1, the thrust measurement test method of the nozzle based on the integrated design of the afterbody and the nozzle of the present embodiment includes the following steps:
s10, designing and processing a wind tunnel test model;
the wind tunnel test model comprises a tail support force measurement test model and a head support spray pipe thrust measurement test model, wherein the model scale is consistent, and the shape of the rear body is consistent; the measuring balance of the tail support force measuring test model is a rod balance 203; the measuring balance of the head support spray pipe thrust measuring test model is a ring balance 4, and the head support spray pipe thrust measuring test model is also provided with a spray pipe 6 and a spray pipe block 21 for replacing the spray pipe 6; the distance between the ring balance 4 of the head support nozzle thrust measurement test model and the rod balance 203 calibration center of the tail support force measurement test model is as follows: the distance delta l in the X direction, the distance delta Y in the Y direction and the distance delta Z in the Z direction;
s20, carrying out a tail support force measurement wind tunnel test;
starting the wind tunnel, and measuring six-component aerodynamic force and aerodynamic moment F of the tail support force measurement test model through the rod balance 203 under the conditions of set incoming flow Mach number and Reynolds number x 、F y 、F z 、m x 、m y 、m z ；
Wherein, F x Axial force for tail support, F y Normal force for tail support, F z Lateral force for tail support, m x Roll moment for tail support, m y Yaw moment, m, for tail support z A pitching moment for the tail support;
s30, performing a measurement test that the head support spray pipe has thrust and outflow;
a nozzle 6 is arranged on the head support nozzle thrust measurement test model; opening the wind tunnel, under the same incoming flow Mach number and Reynolds number conditions as those in the step S20, delivering high-pressure airflow to the spray pipe 6 by the external high-pressure air source, spraying the high-pressure airflow through the spray pipe 6 to generate thrust, and measuring six-component aerodynamic force and aerodynamic moment F of the afterbody under the condition that the thrust has outflow by the ring balance 4 x0 、F y0 、F z0 、m x0 、m y0 、m z0 ；
Wherein, F x0 For axial forces of the afterbody with thrust and outflow, F y0 Normal force of the rear body with thrust and outflow, F z0 For afterbody under the condition of thrust and outflowLateral force of (m) x0 Roll moment of the rear body under the condition of thrust and outflow, m y0 Yaw moment m of the afterbody under the condition of thrust and outflow z0 The pitching moment of the rear body under the condition of thrust and outflow;
s40, performing a measurement test on whether the thrust of the head support spray pipe flows outwards or not;
reserving a nozzle 6 on the head support nozzle thrust measurement test model; closing the wind tunnel, delivering the same high-pressure airflow to the spray pipe 6 by the external high-pressure air source as the high-pressure airflow in the step S20, spraying the high-pressure airflow through the spray pipe 6 to generate thrust, and measuring six-component aerodynamic force and aerodynamic moment F of the afterbody by the ring balance 4 under the condition that the thrust is out-flowing or not x01 、F y01 、F z01 、m x01 、m y01 、m z01 ；
Wherein, F x01 Axial force of the rear body with or without thrust and outflow, F y01 Normal force of the rear body with thrust and without outflow, F z01 M is the lateral force of the rear body under the condition of thrust and no outflow x01 Roll moment m of the rear body under the conditions of thrust and no outflow y01 The yawing moment m of the afterbody under the condition of thrust and no outflow z01 Pitching moment of the afterbody under the condition of thrust and no outflow;
s50, performing a head support spray pipe non-thrust and outflow measurement test;
replacing the nozzle 6 of the head-supported nozzle thrust measurement test model with a nozzle block 21; opening the wind tunnel, and under the same incoming flow Mach number and Reynolds number conditions as those in step S20, measuring the six-component aerodynamic force and aerodynamic moment F of the afterbody under the conditions of existence of thrust and outflow by the ring balance 4 x1 、F y1 、F z1 、m x1 、m y1 、m z1 ；
Wherein, F x1 Axial force of the afterbody in the absence of thrust and with outflow, F y1 Normal force of the rear body in the absence of thrust and with outflow, F z1 The lateral force m of the rear body under the condition of no thrust and outflow x1 Roll moment m of the rear body in the absence of thrust and with outflow y1 Yaw moment m of the afterbody under the condition of no thrust and outflow z1 The pitching moment of the rear body under the condition of no thrust and outflow;
s60, calculating the effective thrust characteristic of the test model in a jet flow state;
calculating the effective thrust characteristic of the test model in the jet flow state according to the test data obtained in the steps S30 and S50:
effective aerodynamic force: f x is effective = F x0 - F x1 ，F y is effective = F y0 - F y1 ，F z is effective = F z0 - F z1 ；
Effective aerodynamic moment: m is x is effective =m x0 -m x1 ，m y is effective =m y0 -m y1 ，m z effective =m z0 -m z1 ；
Wherein, F x is effective Effective axial thrust of the rear body, F y is effective Effective normal thrust of the rear body, F z is effective Effective side thrust of the rear body, m x is effective Effective roll moment of the rear body, m y is effective Effective yaw moment, m, of the rear body z is effective Effective pitching moment of the rear body;
s70, calculating the interference aerodynamic characteristics of the outflow to the thrust characteristics;
calculating the disturbance aerodynamic characteristics of the outflow to the thrust characteristics according to the test data obtained in the steps S30 and S40:
outflow disturbance aerodynamic force: f x interference = F x0 - F x01 ，F y interference = F y0 - F y01 ，F z interference = F z0 - F z01 ；
Outflow disturbance aerodynamic moment: m is x interference =m x0 -m x01 ，m y interference =m y0 -m y01 ，m z interference =m z0 -m z01 ；
Wherein, F x interference Axial force for disturbance of outflow to jet flow, F y interference Normal force for disturbance of outflow to jet flow, F z interference Lateral force for disturbance of outflow to jet flow, m x interference Roll force for disturbance of outflow to jet flowMoment, m y interference Yawing moment m for disturbance of outflow on jet flow z interference The external flow is disturbed to the pitching moment of the jet flow;
s80, calculating the pneumatic characteristic of the full-computer model with the jet flow state;
and (5) correcting and correcting the tail support force measurement test data in the step (S10) by adopting the calculation result in the step (S60) to obtain the pneumatic characteristic of the jet flow state of the full-engine model:
jet state aerodynamic force: f x ’= F x + F x is effective ，F y ’= F y + F y is effective ，F z ’= F z + F z is effective ；
Jet state aerodynamic moment: m is x ’=m x +m x is effective + F y is effective ×Δz+ F z is effective ×Δy，m y ’=m y +m y is effective + F x is effective ×Δz+ F z effective ×Δl，m z ’=m z +m z is effective + F x is effective ×Δy+ F y is effective ×Δl；
Wherein, F x ' axial force in full-machine-band jet state, F y ' Normal force with jet flow state for the whole machine, F z ' lateral force in full machine with jet flow state, m x ' Rolling moment m in the state of full machine with jet flow y ' yaw moment m in a full-machine band jet state z ' is the pitching moment of the whole machine with the jet flow state.
As shown in FIG. 2, the tail-supported force-measuring test model comprises a host model 201 provided with a tail nozzle 202, a rod balance 203 is arranged in an inner cavity of the host model 201, the rear end of the rod balance 203 is connected with a tail support rod 204, and the rear end of the tail support rod 204 is fixed on a wind tunnel middle support.
As shown in fig. 3~9, the head support nozzle thrust measurement test model is installed at the outlet of the wind tunnel nozzle of the high-speed free jet wind tunnel, the ventilating blade 2, the ventilating support rod 3, the ring balance 4, the aircraft rear body 5, the corrugated pipe 20, the nozzle 6 and the fairing 7 are located in the wind tunnel test section, wherein the space outside the ventilating blade 2 is a plenum chamber 1 of the high-speed free jet wind tunnel, and an airflow pipeline of a high-pressure air source enters the plenum chamber 1 through a high-pressure air pipe inlet 19 of the plenum chamber 1 and is fixedly connected with an interface of the blade airflow channel 16.
The head support nozzle thrust measurement test model comprises a ventilating blade 2, a ventilating support rod 3, a ring balance 4, an aircraft rear body 5, a nozzle 6, a fairing 7 and a corrugated pipe 20; also comprises a nozzle block 21 for replacing the nozzle 6;
the ventilation support rod 3 is a circular tube body I, the front end of the ventilation support rod 3 is closed, and the rear end of the ventilation support rod is opened;
a fairing 7 is arranged on the end head of the front end of the ventilating support rod 3; the circumferential direction of the front end of the ventilation support rod 3 is provided with ventilation blades 2 which are axially symmetrically distributed;
the rear end of the ventilating support rod 3 is sequentially connected with a measuring section, an aircraft rear body 5 and a spray pipe 6, the aircraft rear body 5 is a circular pipe body II, and the spray pipe 6 is an axisymmetric spray pipe; the measuring section comprises a ring balance 4 and a corrugated pipe 20 which are sleeved in sequence from outside to inside, the fixed end of the ring balance 4 is arranged on the outer layer installation surface at the rear end of the ventilating support rod 3, the free end of the ring balance 4 is arranged on the outer layer installation surface at the front end of the rear body 5 of the aircraft, the front end of the corrugated pipe 20 is arranged on the inner layer installation surface at the rear end of the ventilating support rod 3, and the rear end of the corrugated pipe 20 is arranged on the inner layer installation surface at the front end of the rear body 5 of the aircraft;
the central axes of the fairing 7, the ventilating blade 2, the ventilating support rod 3, the ring balance 4, the aircraft rear body 5 and the corrugated pipe 20 are coincided with the central axis of the wind tunnel test section; an included angle alpha is formed between the central axis of the spray pipe 6 and the central axis of the wind tunnel test section, and alpha is a spray pipe deflection angle;
a plurality of blade airflow channels 16 communicated with an external high-pressure air source are arranged inside the ventilation blade 2, airflow outlets 17 of the blade airflow channels 16 are positioned on the inner wall of the ventilation support rod 3, and a ventilation support rod airflow channel 18 is arranged on the central axis of the ventilation support rod 3; high-pressure air source airflow enters the ventilating support rod airflow channel 18 along the blade airflow channel 16 through the airflow outlet 17 and then is sprayed out through the corrugated pipe 20 and the spray pipe 6; the ring balance 4 measures aerodynamic force and aerodynamic moment of the spray pipe 6; the bellows 20 deforms along with the ring balance 4 and seals and isolates the high-pressure air source airflow in the ventilation strut airflow channel 18, so that the high-pressure air source airflow is prevented from affecting the measurement of the ring balance 4.
The deflection angle alpha of the spray pipe ranges from minus 20 degrees to 20 degrees.
The windward side of the ventilating vane 2 is provided with an arc windward side for rectification.
The front end of the ventilation support rod 3 is sealed by a blocking cover 8; the blocking cover 8 is a step cylinder, and the diameter of the front section cylinder of the blocking cover 8 is smaller than that of the rear section cylinder; the rear section of the cylinder of the blocking cover 8 is inserted into the front end of the ventilation support rod 3; the front section cylinder of the blocking cover 8 extends out of the ventilating support rod 3, and the cutting sleeve 10 is sleeved on the front section cylinder of the blocking cover 8; the rear end of the cutting sleeve 10 is provided with a circular groove matched with the front end head of the ventilation support rod 3, the rear end face of the cutting sleeve 10 is a contact surface of the cutting sleeve 10 and the front end face of the ventilation blade 2, and the front end of the cutting sleeve 10 is provided with a locking nut 9; the locking nut 9 is screwed down, the clamping sleeve 10 props against the front end face of the ventilating vane 2, and the blocking cover 8 and the clamping sleeve 10 are fixed;
the fairing 7 is sleeved on the clamping sleeve 10 and wraps the blocking cover 8, the locking nut 9 and the clamping sleeve 10, and the spherical outer surface of the fairing 7 faces the incoming flow and is used for incoming flow rectification.
The inner wall of the cutting sleeve 10 is provided with a sealing ring I11 for sealing the contact surface of the cutting sleeve 10 and the front section cylinder of the blocking cover 8; the rear end face of the cutting sleeve 10 is provided with a sealing ring II 12 which is used for sealing the contact surface between the rear end face of the cutting sleeve 10 and the front end face of the ventilation blade 2; a sealing ring III 13 is arranged on the rear section of the cylinder of the blocking cover 8 and used for sealing the contact surface of the blocking cover 8 and the inner wall of the ventilating support rod 3; a sealing ring IV 14 is arranged at the front section of the ventilation support rod 3 and used for sealing the contact surface of the ventilation support rod 3 and the rear end surface of the ventilation blade 2; the front end of the corrugated pipe 20 is provided with a sealing ring V15 for sealing the contact surface between the inner layer installation surface at the rear end of the ventilation strut 3 and the front end of the corrugated pipe 20, and the front end of the corrugated pipe 20 is also provided with a sealing ring V15 for sealing the contact surface between the inner layer installation surface at the front end of the aircraft rear body 5 and the rear end of the corrugated pipe 20.
Although the embodiments of the present invention have been disclosed above, it is not limited to the applications listed in the description and the embodiments, and it can be fully applied to various fields of high-speed free jet pressure matching pattern methods suitable for the present invention. Additional modifications and refinements will readily occur to those skilled in the art without departing from the principles of the present invention, and the present invention is not limited to the specific details and illustrations shown and described herein.

Claims (4)

1. A thrust measurement test method of a spray pipe based on the integrated design of a afterbody and the spray pipe is characterized by comprising the following steps of:
s10, designing and processing a wind tunnel test model;
the wind tunnel test model comprises a tail support force measurement test model and a head support spray pipe thrust measurement test model, wherein the model scale is consistent, and the shape of the rear body is consistent;
the tail support force measurement test model comprises a host model (201) provided with a tail nozzle (202), a rod balance (203) is arranged in an inner cavity of the host model (201), the rear end of the rod balance (203) is connected with a tail support rod (204), and the rear end of the tail support rod (204) is fixed on a wind tunnel middle support;
the head support spray pipe thrust measurement test model comprises ventilation blades (2), a ventilation support rod (3), a ring balance (4), an aircraft rear body (5), a spray pipe (6), a fairing (7) and a corrugated pipe (20); also comprises a nozzle block (21) for replacing the nozzle (6);
the ventilation support rod (3) is a circular tube body I, the front end of the ventilation support rod (3) is closed, and the rear end of the ventilation support rod is opened;
a fairing (7) is arranged on the end head of the front end of the ventilating support rod (3); the circumferential direction of the front end of the ventilating support rod (3) is provided with ventilating vanes (2) which are axially symmetrically distributed;
the rear end of the ventilating support rod (3) is sequentially connected with the measuring section, the aircraft rear body (5) and the spray pipe (6), the aircraft rear body (5) is a circular pipe body II, and the spray pipe (6) is an axisymmetric spray pipe; the measuring section comprises a ring balance (4) and a corrugated pipe (20) which are sleeved in sequence from outside to inside, the fixed end of the ring balance (4) is installed on the outer layer installation surface of the rear end of the ventilating support rod (3), the free end of the ring balance (4) is installed on the outer layer installation surface of the front end of the rear body (5) of the aircraft, the front end of the corrugated pipe (20) is installed on the inner layer installation surface of the rear end of the ventilating support rod (3), and the rear end of the corrugated pipe (20) is installed on the inner layer installation surface of the front end of the rear body (5) of the aircraft;
the central axes of the fairing (7), the ventilating blade (2), the ventilating support rod (3), the ring balance (4), the aircraft rear body (5) and the corrugated pipe (20) are superposed with the central axis of the wind tunnel test section; an included angle alpha is formed between the central axis of the spray pipe (6) and the central axis of the wind tunnel test section, and alpha is a spray pipe deflection angle;
a plurality of blade airflow channels (16) communicated with an external high-pressure air source are arranged in the ventilating blade (2), airflow outlets (17) of the blade airflow channels (16) are positioned on the inner wall of the ventilating support rod (3), and a ventilating support rod airflow channel (18) is arranged on the central axis of the ventilating support rod (3); airflow of a high-pressure air source enters an airflow channel (18) of the ventilating support rod through an airflow outlet (17) along an airflow channel (16) of the blade and then is sprayed out through the corrugated pipe (20) and the spray pipe (6); the ring balance (4) measures the aerodynamic force and aerodynamic moment of the spray pipe (6); the corrugated pipe (20) deforms along with the ring balance (4), and seals and isolates the high-pressure air source airflow in the airflow channel (18) of the ventilating support rod, so that the high-pressure air source airflow is prevented from influencing the measurement of the ring balance (4);
the front end of the ventilation support rod (3) is sealed by a blocking cover (8); the blocking cover (8) is a step cylinder, and the diameter of the front section cylinder of the blocking cover (8) is smaller than that of the rear section cylinder; the rear section of the cylinder of the blocking cover (8) is inserted into the front end of the ventilation support rod (3); the front section cylinder of the blocking cover (8) extends out of the ventilating support rod (3), and the cutting sleeve (10) is sleeved on the front section cylinder of the blocking cover (8); the rear end of the cutting sleeve (10) is provided with a circular groove matched with the front end head of the ventilation support rod (3), the rear end face of the cutting sleeve (10) is a contact face of the cutting sleeve (10) and the front end face of the ventilation blade (2), and the front end of the cutting sleeve (10) is provided with a locking nut (9); the locking nut (9) is screwed down, the clamping sleeve (10) is tightly pressed against the front end face of the ventilating blade (2), and the plugging cover (8) and the clamping sleeve (10) are fixed;
the fairing (7) is sleeved on the clamping sleeve (10) and wraps the blocking cover (8), the locking nut (9) and the clamping sleeve (10), and the spherical outer surface of the fairing (7) faces the incoming flow and is used for incoming flow rectification;
the measuring balance of the tail support force measuring test model is a rod balance (203); the measuring balance of the head-supported spray pipe thrust measuring test model is a ring balance (4), and the head-supported spray pipe thrust measuring test model is also provided with a spray pipe (6) and a spray pipe blocking block (21) for replacing the spray pipe (6); the distance between a ring balance (4) of the head support nozzle thrust measurement test model and a rod balance (203) calibration center of the tail support force measurement test model is as follows: the distance delta l in the X direction, the distance delta Y in the Y direction and the distance delta Z in the Z direction;
s20, carrying out a tail support force measurement wind tunnel test;
starting the wind tunnel, and measuring six-component aerodynamic force and aerodynamic moment F of the tail support force measurement test model by the rod balance (203) under the conditions of set incoming flow Mach number and Reynolds number x 、F y 、F z 、m x 、m y 、m z ；
Wherein, F x Axial force for tail support, F y Normal force for tail support, F z Lateral force for tail support, m x Roll moment for tail support, m y Yaw moment m for tail support z A pitching moment for the tail support;
s30, performing a measurement test that the head support spray pipe has thrust and outflow;
a nozzle (6) is arranged on the head support nozzle thrust measurement test model; opening the wind tunnel, under the same incoming flow Mach number and Reynolds number conditions as those in the step S20, delivering high-pressure airflow to the spray pipe (6) by the external high-pressure air source, spraying the high-pressure airflow out of the spray pipe (6) to generate thrust, and measuring six-component aerodynamic force and aerodynamic moment F of the afterbody under the condition that the thrust has outflow by the ring balance (4) x0 、F y0 、F z0 、m x0 、m y0 、m z0 ；
Wherein, F x0 For axial forces of the afterbody with thrust and outflow, F y0 Normal force of the rear body under thrust and outflow conditions, F z0 For the side force of the rear body with thrust and outflow, m x0 Roll moment of the rear body under thrust and outflow conditions, m y0 For yawing moment m of the afterbody under the conditions of thrust and outflow z0 The pitching moment of the rear body under the condition of thrust and outflow;
s40, performing a measurement test on whether the thrust of the head support spray pipe flows outwards or not;
reserving a nozzle (6) on the head support nozzle thrust measurement test model; closing the wind tunnel, delivering the same high-pressure airflow as the high-pressure airflow in the step S20 to the spray pipe (6) by an external high-pressure air source, spraying the high-pressure airflow out of the spray pipe (6) to generate thrust, and measuring six-component aerodynamic force and aerodynamic moment F of the afterbody under the condition that the thrust is out-flowing or not by the ring balance (4) x01 、F y01 、F z01 、m x01 、m y01 、m z01 ；
Wherein, F x01 Axial force of the rear body with or without thrust and outflow, F y01 Normal force of the rear body with thrust and without outflow, F z01 M is the lateral force of the rear body under the condition of thrust and no outflow x01 Roll moment m of the rear body under the condition of thrust and no outflow y01 The yawing moment m of the afterbody under the condition of thrust and no outflow z01 Pitching moment of the rear body under the condition of thrust and no outflow;
s50, performing a non-thrust and outflow measurement test on the head support spray pipe;
replacing a nozzle (6) of the head-supported nozzle thrust measurement test model with a nozzle block (21); opening the wind tunnel, and under the same incoming flow Mach number and Reynolds number conditions as the step S20, measuring six-component aerodynamic force and aerodynamic moment F of the afterbody under the conditions of no thrust and outflow by the ring balance (4) x1 、F y1 、F z1 、m x1 、m y1 、m z1 ；
Wherein, F x1 Axial force of the rear body in the absence of thrust and with outflow, F y1 Normal force of the rear body in the absence of thrust and with outflow, F z1 The lateral force m of the rear body under the condition of no thrust and outflow x1 Roll moment m of the rear body in the absence of thrust and with outflow y1 Yaw moment m of the afterbody under the condition of no thrust and outflow z1 The pitching moment of the rear body under the condition of no thrust and outflow;
s60, calculating the effective thrust characteristic of the test model in a jet flow state;
calculating the effective thrust characteristic of the test model in the jet flow state according to the test data obtained in the steps S30 and S50:
effective aerodynamic force: f x is effective = F x0 - F x1 ，F y is effective = F y0 - F y1 ，F z is effective = F z0 - F z1 ；
Effective aerodynamic moment: m is x is effective =m x0 -m x1 ，m y is effective =m y0 -m y1 ，m z is effective =m z0 -m z1 ；
Wherein, F x is effective Effective axial thrust of the rear body, F y is effective Effective normal thrust of the rear body, F z is effective Effective side thrust of the rear body, m x is effective Effective roll moment of the rear body, m y is effective Effective yaw moment, m, of the rear body z effective Effective pitching moment of the rear body;
s70, calculating the interference aerodynamic characteristics of the outflow flow on the thrust characteristics;
calculating the disturbance aerodynamic characteristics of the outflow to the thrust characteristics according to the test data obtained in the steps S30 and S40:
outflow disturbance aerodynamic force: f x interference = F x0 - F x01 ，F y interference = F y0 - F y01 ，F z interference = F z0 - F z01 ；
Outflow disturbance aerodynamic moment: m is x interference =m x0 -m x01 ，m y interference =m y0 -m y01 ，m z interference =m z0 -m z01 ；
Wherein, F x interference Axial force for disturbance of outflow to jet flow, F y interference Normal force for disturbance of outflow to jet flow, F z interference Lateral force for disturbance of outflow to jet flow, m x interference Roll moment for disturbance of outflow to jet flow, m y interference Yawing moment m for disturbance of outflow on jet flow z interference The external flow is disturbed to the pitching moment of the jet flow;
s80, calculating the pneumatic characteristic of the full-computer model with a jet flow state;
and (5) correcting and correcting the tail support force measurement test data in the step (S10) by adopting the calculation result in the step (S60) to obtain the pneumatic characteristic of the jet flow state of the full-engine model:
jet state aerodynamic force: f x ’= F x + F x is effective ，F y ’= F y + F y is effective ，F z ’= F z + F z is effective ；
Jet state aerodynamic moment: m is x ’=m x +m x is effective + F y is effective ×Δz+ F z is effective ×Δy，m y ’=m y +m y is effective + F x is effective ×Δz+ F z is effective ×Δl，m z ’=m z +m z is effective + F x is effective ×Δy+ F y is effective ×Δl；
Wherein, F x ' axial force in the form of a full-machine jet, F y ' Normal force with jet flow state for the whole machine, F z ' lateral force in full machine with jet flow state, m x ' Rolling moment m in the state of full machine with jet flow y ' yaw moment m in a full-machine band jet state z ' is the pitching moment of the whole machine with the jet flow state.
2. The nozzle thrust measurement test method based on the afterbody and nozzle integrated design of claim 1, wherein the nozzle deflection angle α is in the range of-20 °.
3. The nozzle thrust measurement test method based on the afterbody and nozzle integrated design of claim 2, characterized in that the windward side of the ventilating vane (2) is provided with an arc windward side for rectification.
4. The nozzle thrust measurement test method based on the afterbody and nozzle integrated design of claim 3, characterized in that the inner wall of the cutting ferrule (10) is provided with a sealing ring I (11) for sealing the contact surface of the cutting ferrule (10) and the front section cylinder of the blanking cap (8); the rear end face of the cutting sleeve (10) is provided with a sealing ring II (12) for sealing the contact surface of the rear end face of the cutting sleeve (10) and the front end face of the ventilating blade (2); a sealing ring III (13) is arranged on the rear section of the cylinder of the blocking cover (8) and used for sealing the contact surface of the blocking cover (8) and the inner wall of the ventilating support rod (3); a sealing ring IV (14) is arranged at the front section of the ventilation support rod (3) and is used for sealing the contact surface of the ventilation support rod (3) and the rear end surface of the ventilation blade (2); the front end of the corrugated pipe (20) is provided with a sealing ring V (15) for sealing the contact surface between the inner layer mounting surface at the rear end of the ventilation support rod (3) and the front end of the corrugated pipe (20), and the front end of the corrugated pipe (20) is also provided with the sealing ring V (15) for sealing the contact surface between the inner layer mounting surface at the front end of the aircraft rear body (5) and the rear end of the corrugated pipe (20).
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