Equivalent state wind tunnel test method of hypersonic oblique detonation engine

Abstract

The application provides a hypersonic oblique detonation engine equivalent state wind tunnel experiment method, which mainly comprises the following steps: (1) obtaining the inlet parameters of the combustion chamber of the oblique detonation experiment model; (2) determining the flow field simulation capability of the experiment wind tunnel; (3) calculating the equivalent combustion chamber inlet parameters and determining the equivalent experiment scheme. The application can simulate the combustion chamber inlet parameters of the oblique detonation engine under a high Mach number state by using a low Mach number wind tunnel, thereby expanding the experiment conditions for carrying out the combustion and thrust performance wind tunnel experiment of the oblique detonation engine and promoting the engineering development of the oblique detonation engine.

Description

Technical Field

[0001] This invention relates to the field of oblique detonation engine technology, and specifically to a wind tunnel experimental method for the equivalent state of a hypersonic oblique detonation engine. Background Technology

[0002] The oblique detonation engine is a novel propulsion system for air-breathing hypersonic vehicles. It utilizes oblique detonation waves to achieve efficient combustion in supersonic airflow, offering advantages such as rapid heat release, high specific impulse, short combustion chamber, and fewer moving parts. Thanks to the kilometer-per-second propagation speed of the detonation wave, the inlet airflow of the oblique detonation engine combustion chamber can maintain a high velocity. Strong shock wave ignition can achieve the conversion of fuel chemical energy into internal energy within millimeters, significantly reducing the combustion chamber size and possessing potential engineering advantages. It represents a cutting-edge research area in hypersonic propulsion technology. Theoretically, the thrust performance of engines based on detonation combustion can be more than 30% higher than that of existing scramjet engines based on isobaric combustion. Through effective combustion organization, entropy increase and total pressure loss can be further reduced, thereby further optimizing the engine's thrust performance. It is considered a more suitable air-breathing propulsion method for hypersonic vehicles above Mach 9, capable of breaking through the current Mach number limit of hydrocarbon fuel scramjet engines. Due to the highly complex physical and chemical processes occurring in the combustion chamber of a slant-detonation engine, in-depth research into the slant-detonation combustion mechanism, including its initiation characteristics, unsteady properties, and wavefront stability, is crucial for the engineering development of slant-detonation engines. While numerical simulation studies of slant-detonation combustion phenomena and mechanisms are relatively advanced, the engineering development of slant-detonation engines is still in its early stages. Wind tunnel testing of slant-detonation engine combustion performance and components is urgently needed to advance the integrated design of the entire engine. The main problem with conducting wind tunnel tests on slant-detonation engines is that these engines operate at Mach numbers greater than 9. Conventional wind tunnel testing methods use the same Mach number in the wind tunnel flow field as the engine's flight Mach number, placing the engine model within the flow field for testing. The main difficulties in conducting wind tunnel experiments on oblique detonation engines using this approach are as follows: Because the oblique detonation chemical reaction process is closely related to the incoming flow temperature, pressure, and scale of the model, to accurately reflect the chemical reaction process during flight, the total temperature and pressure of the incoming flow and the scale of the wind tunnel model must be identical to those of the actual aircraft. This results in high total temperature and pressure of the incoming flow and a large flow field scale in the wind tunnel. Furthermore, when using aviation kerosene as fuel, the increased fuel ignition delay requires a longer effective wind tunnel testing time. Currently, the number of wind tunnels with Mach numbers above 9 that can meet the experimental conditions for oblique detonation engines is limited, which is detrimental to the development and testing of oblique detonation engines. Summary of the Invention

[0003] To address the technical problems existing in the background art, this invention proposes an equivalent state wind tunnel test method for hypersonic oblique detonation engines. This method can use a low Mach number wind tunnel to simulate the combustion chamber inlet parameters of an oblique detonation engine under high Mach number conditions, thereby broadening the experimental conditions for conducting wind tunnel tests on the combustion and thrust performance of oblique detonation engines and promoting the engineering development of oblique detonation engines.

[0004] To solve the above-mentioned technical problems, the present invention provides a wind tunnel experimental method for the equivalent state of a hypersonic oblique detonation engine, which mainly includes the following steps:

[0005] (1) Obtain the inlet parameters of the combustion chamber of the oblique detonation experimental model;

[0006] (2) Determine the flow field simulation capability of the experimental wind tunnel;

[0007] (3) Calculate the inlet parameters of the equivalent combustion chamber and determine the equivalent experimental scheme.

[0008] The wind tunnel test method for the equivalent state of the hypersonic oblique detonation engine, wherein the specific process of step (1) is as follows:

[0009] (1.1) Define the wind tunnel test mission for the oblique detonation engine and determine the flight Mach number Ma of the oblique detonation test model according to the requirements of the test mission. fly Flight status, high-altitude static temperature (T) fly Static pressure P fly and the compression angles at each stage of the intake manifold;

[0010] (1.2) The Mach number at the combustor inlet during flight was calculated using the method of calculating the aftershock parameters of the oblique shock wave. ode Flight state combustion chamber inlet static temperature T ode and the static pressure P at the combustion chamber inlet during flight ode .

[0011] The wind tunnel test method for the equivalent state of the hypersonic oblique detonation engine, wherein step (2) specifically includes the following steps:

[0012] (2.1) Determine the Mach number of the wind tunnel nozzle. nozzle The range of variation;

[0013] (2.2) Determine the range of variation of total temperature T0 and total pressure P0 in the wind tunnel flow field.

[0014] The wind tunnel test method for the equivalent state of the hypersonic oblique detonation engine, wherein the specific process of step (3) is as follows:

[0015] (3.1) Determine the allowable percentage differences in static temperature, static pressure, and Mach number between the equivalent state and the flight state, T%, P%, and Ma%, and give the compression angle settings for each stage of the equivalent inlet;

[0016] (3.2) Given the total temperature, total pressure, and Mach number of the flow field;

[0017] (3.3) Calculate the static temperature T of the free flow field at the nozzle exit based on the isentropic expansion relationship. flow Static pressure P flow ;

[0018] (3.4) Based on the given first-stage compression angle θ1 of the inlet and the calculated static temperature T of the free flow field at the nozzle exit flow Static pressure P flow The flow field parameters after the first stage compression in the inlet were calculated using the oblique shock wave post-parameter calculation method.

[0019] (3.5) Repeat step (3.4) to complete the parameter calculation after compression of each stage of the intake, and obtain the equivalent parameters of the combustion chamber inlet of the oblique detonation engine: equivalent state combustion chamber inlet static temperature T equi Equivalent state combustion chamber inlet static pressure P equi and equivalent state combustion chamber inlet Mach number Ma equi ;

[0020] (3.6) Calculate whether the differences in static temperature, static pressure, and Mach number at the combustion chamber inlet under equivalent and flight conditions satisfy the following three conditions:

[0021]

[0022] (3.7) If the three conditions in step (3.6) above are met, stop the calculation; otherwise, change the total temperature, total pressure and Mach number of the free flow field at the nozzle exit, and repeat steps (3.2)-(3.6) above to complete the calculation under all experimental wind tunnel flow field simulation capability conditions.

[0023] The wind tunnel test method for the equivalent state of the hypersonic oblique detonation engine, wherein the specific process of step (3.3) is as follows:

[0024] (3.3.1) The wind tunnel accelerates stationary gas through isentropic expansion via a nozzle, forming a free flow field with a certain velocity at the nozzle exit. The stationary gas is located in the stagnation chamber of the wind tunnel, and the entropy value remains constant during the isentropic expansion process. Given the total temperature P0, total pressure T0, and nozzle exit Mach number Ma of the flow field. nozzle ;

[0025] (3.3.2) Calculate the entropy S0 and enthalpy H0 of the resident gas;

[0026]

[0027] In the formula, R is the gas constant, and c1, c2, c3, c4, c5, b1, and b2 are piecewise constants with respect to temperature;

[0028] (3.3.3) Set the pressure expansion rate dP during the isentropic expansion process;

[0029] (3.3.4) The isentropic expansion process is calculated in multiple iterative steps, with the pressure change at each step being P. i+1 =P i -dP, due to isentropic expansion, the entropy S remains constant; calculate the local gas enthalpy H. i+1 and speed of sound a i+1 :

[0030]

[0031] In the formula, γ is the specific heat ratio of the gas;

[0032] Due to energy conservation during nozzle expansion, the local gas velocity u i+1 :

[0033] u i+1 =sqrt(H i +(u i ) 2 / 2.0-H i+1 );

[0034] (3.3.5) The local gas Mach number is:

[0035] Ma i+1 =u i+1 / a i+1 ;

[0036] (3.3.6) Compare whether the gas Mach number reaches the nozzle Mach number. nozzle If the target is not reached, repeat steps (3.3.4)-(3.3.5) until the nozzle exit Mach number is reached. nozzle ;

[0037] (3.3.7) Based on the pressure P of the last iteration step i+1 The static temperature T of the free-flowing gas at the nozzle exit is calculated from the entropy value S. flow Static pressure P flow .

[0038] The wind tunnel test method for the equivalent state of the hypersonic oblique detonation engine, wherein the calculation method for the oblique shock wave parameters in steps (1.2) and (3.4) is as follows:

[0039] The forward and backward tangential and normal parameters of the oblique shock wave satisfy the following relationship:

[0040] w1 = u1sinβ;

[0041] w2 = u2sin(β-θ);

[0042] v1 = u1cosβ;

[0043] v2 = u2cos(β-θ);

[0044] Where β is the oblique shock wave angle, θ is the inlet compression angle, w and v are the normal and tangential velocity components of velocity u, respectively, subscript 1 indicates the parameters before the oblique shock wave, and subscript 2 indicates the parameters after the oblique shock wave;

[0045] Since v1 = v2, then:

[0046] w2 = u1cosβtan(β-θ);

[0047] From the above relationships, the shock angle and compression angle can be expressed as functions of the incoming tangential velocity and normal velocity, i.e.:

[0048] β=sin -1 (w1 / u1);

[0049]

[0050] The parameters before and after the shock wave simultaneously satisfy the following conservation relationships of mass, momentum, and energy:

[0051] ρ1w1=ρ2w2;

[0052]

[0053] In the formula, ρ, P, and H represent the gas density, pressure, and enthalpy, respectively.

[0054] Based on the above relationships between the tangential and normal parameters before and after the oblique shock wave, as well as the conservation relationships of mass, momentum, and energy, the parameters after the oblique shock wave are calculated. The specific process is as follows:

[0055] (1.2.01) Given the incoming flow parameters T1, P1, u1 and compression angle θ of the oblique shock wave;

[0056] (1.2.02) Given the initial value of the oblique shock wave angle β, calculate the normal velocity component w1 before the oblique shock wave;

[0057] (1.2.03) Given the initial value of the back density ρ2 of the oblique shock wave, calculate the specific volume and back pressure:

[0058]

[0059] In the formula, V is the specific volume, P is the pressure, and ρ is the density;

[0060] (1.2.04) Calculate the normal velocity after the oblique shock wave and the difference function of momentum and energy before and after the oblique shock wave:

[0061]

[0062] In the formula, P and H are the difference functions of momentum and energy before and after the oblique shock wave, respectively;

[0063] (1.2.05) Keeping the specific volume constant, calculate H(T2+ΔT) and P(T2+ΔT) under the condition of temperature disturbance ΔT;

[0064] (1.2.06) Keeping the temperature constant, calculate H(V2+ΔV) and P(V2+ΔV) under specific volume disturbance ΔV;

[0065] (1.2.07) Calculate the Jacobian matrix using first-order difference:

[0066]

[0067] (1.2.08) Calculate the system of linear equations Determine the temperature and specific volume corrections δT and δV;

[0068] (1.2.09) Restrictions are imposed on δT and δV: when |δT|>0.2*T², then 0.2*T²*sgn(δT); when |δV|>0.2*V², then 0.2*T²*sgn(δT) is taken; and when |δV|>0.2*V², then 0.2*T²*sgn(δT) is taken. The value is 0.2*V2*sgn(δV) when |δV|>0.5*(V1-V2) and when V2+δV>V1, the value is 0.5*(V1-V2)*sgn(δV), where sgn is the sign function;

[0069] (1.2.10) Determine the new oblique shock wave back temperature T′2 and specific volume V′2:

[0070] T′2=T2-δT;

[0071] V′2=V2-δV;

[0072] (1.2.11) Check if the convergence condition has been met:

[0073] T′2-T2<T Err ;

[0074] V′2-V2<V Err ;

[0075] In the formula, T Err V Err Allowable errors for the set temperature and specific volume;

[0076] (1.2.12) Repeat steps (1.2.03) to (1.2.11) until the convergence condition is met;

[0077] (1.2.13) Determine the pressure, velocity and Mach number of the oblique shock wave based on the temperature and specific volume behind the oblique shock wave.

[0078] By adopting the above technical solution, the present invention has the following beneficial effects:

[0079] The present invention provides a wind tunnel test method for the equivalent state of a hypersonic oblique detonation engine. This method can use a low Mach number wind tunnel to simulate the combustion chamber inlet parameters of an oblique detonation engine under high Mach number conditions, thereby broadening the experimental conditions for conducting wind tunnel tests on the combustion and thrust performance of oblique detonation engines. Compared with conventional methods, it can reduce the wind tunnel drive capability requirements, expand the test range of oblique detonation engines, and promote the engineering development of oblique detonation engines. Attached Figure Description

[0080] To more clearly illustrate the specific embodiments of the present invention or the technical solutions in the prior art, the accompanying drawings used in the description of the specific embodiments or the prior art will be briefly introduced below. Obviously, the drawings described below are some embodiments of the present invention. For those skilled in the art, other drawings can be obtained based on these drawings without creative effort.

[0081] Figure 1 This is a flowchart of the wind tunnel test method for the equivalent state of the hypersonic oblique detonation engine of the present invention;

[0082] Figure 2 This diagram illustrates the relationship between the tangential and normal parameters of the oblique shock wave involved in the wind tunnel experimental method for the equivalent state of the hypersonic oblique detonation engine of this invention.

[0083] Figure 3 A conventional wind tunnel test scheme for a two-stage compression inlet engine at Mach 10.

[0084] Figure 4 This is an equivalent wind tunnel experimental scheme for conducting experiments using a single-stage compression inlet at Mach number 6. Detailed Implementation

[0085] The technical solution of the present invention will now be clearly and completely described with reference to the accompanying drawings. Obviously, the described embodiments are only some, not all, of the embodiments of the present invention. Based on the embodiments of the present invention, all other embodiments obtained by those skilled in the art without creative effort are within the scope of protection of the present invention.

[0086] The present invention will be further explained below with reference to specific embodiments.

[0087] Conventional wind tunnel testing of hypersonic oblique detonation engines involves placing an engine model within the same flow field at the same Mach number as the engine's flight path. Based on the working principles of wind tunnels and oblique detonation engines, the wind tunnel flow field is formed by the isentropic expansion of the resident gas through the nozzle. An oblique detonation engine mainly consists of three parts: the inlet, the combustion chamber, and the exhaust nozzle. The inlet's primary function is to compress the incoming air, increasing its pressure and temperature to create conditions for oblique detonation combustion within the combustion chamber. Therefore, in conventional oblique detonation wind tunnel experiments, air first expands through the nozzle and then is compressed through the inlet to form the necessary flow conditions for the combustion chamber. Since expansion and compression are opposite physical processes, by reducing the wind tunnel nozzle expansion process and optimizing the inlet compression angle design, it is possible to simulate the combustion chamber inlet parameters of an oblique detonation engine at high Mach numbers using a low-Mach number wind tunnel. This expands the experimental conditions for conducting wind tunnel tests on the combustion and thrust performance of oblique detonation engines.

[0088] like Figure 1 As shown in the figure, the wind tunnel test method for the equivalent state of a hypersonic oblique detonation engine provided in this embodiment specifically includes the following steps:

[0089] (1) Obtain the inlet parameters of the combustion chamber of the oblique detonation experimental model.

[0090] (1.1) Define the wind tunnel test mission for the oblique detonation engine and determine the flight Mach number Ma of the oblique detonation test model according to the requirements of the test mission. fly Flight status, high-altitude static temperature (T) fly Static pressure P fly and the compression angles at each stage of the intake manifold;

[0091] (1.2) The Mach number at the combustor inlet during flight was calculated using the method of calculating the aftershock parameters of the oblique shock wave. ode Flight state combustion chamber inlet static temperature T ode and the static pressure P at the combustion chamber inlet during flight ode ;

[0092] (2) Determine the flow field simulation capability of the experimental wind tunnel

[0093] (2.1) Determine the Mach number of the wind tunnel nozzle. nozzle The range of variation;

[0094] (2.2) Determine the range of variation of total temperature T0 and total pressure P0 in the wind tunnel flow field;

[0095] (3) Calculate the inlet parameters of the equivalent combustion chamber and determine the equivalent experimental scheme.

[0096] (3.1) Determine the allowable percentage differences in static temperature, static pressure, and Mach number between the equivalent state and the flight state, T%, P%, and Ma%, and give the compression angle settings for each stage of the equivalent inlet;

[0097] (3.2) Given the total temperature, total pressure, and Mach number of the flow field,

[0098] (3.3) Calculate the static temperature T of the free flow field at the nozzle exit based on the isentropic expansion relationship. flow Static pressure P flow

[0099] (3.4) Based on the given first-stage compression angle θ1 of the inlet and the calculated static temperature T of the free flow field at the nozzle exit flow Static pressure P flow The flow field parameters after the first stage compression in the inlet were calculated using the oblique shock wave post-parameter calculation method.

[0100] (3.5) Repeat step (3.4) to complete the parameter calculation after compression of each stage of the intake, and obtain the equivalent parameters of the combustion chamber inlet of the oblique detonation engine: equivalent state combustion chamber inlet static temperature T equi Equivalent state combustion chamber inlet static pressure P equi Equivalent state combustion chamber inlet Mach number Ma equi .

[0101] (3.6) Calculate whether the differences between the equivalent state and the flight state combustor inlet static temperature, static pressure, and Mach number meet the following three conditions:

[0102]

[0103] (3.7) If the three conditions in step (3.6) are met, stop the calculation; otherwise, change the total temperature, total pressure and Mach number of the free flow field at the nozzle exit, and repeat steps (3.2)-(3.6) to complete the calculation under all experimental wind tunnel flow field simulation capability conditions.

[0104] The isentropic expansion calculation process in step (3.3) above is as follows:

[0105] (3.3.1) The wind tunnel accelerates stationary gas through isentropic expansion via nozzles to form a free flow field with a certain velocity at the nozzle exit. The stationary gas is in the stagnation chamber of the wind tunnel, and the entropy value remains unchanged during the isentropic expansion process.

[0106] Given the total temperature P0, total pressure T0, and nozzle exit Mach number Ma of the flow field. nozzle ;

[0107] (3.3.2) Calculate the entropy S0 and enthalpy H0 of the resident gas;

[0108]

[0109] In the formula, R is the gas constant, and c1, c2, c3, c4, c5, b1, and b2 are piecewise constants with respect to temperature.

[0110] (3.3.3) Set the pressure expansion rate dP during the isentropic expansion process;

[0111] (3.3.4) The isentropic expansion process is calculated in multiple iterative steps, with the pressure change at each step being P. i+1 =P i -dP, due to isentropic expansion, the entropy S remains constant; calculate the local gas enthalpy H. i+1 and speed of sound a i+1 ,

[0112]

[0113]

[0114] In the formula, γ is the specific heat ratio of the gas.

[0115] Due to energy conservation during nozzle expansion, the local gas velocity u i+1 :

[0116] u i+1 =sqrt(H i +(u i ) 2 / 2.0-H i+1 );

[0117] (3.3.5) The local gas Mach number is:

[0118] Ma i+1 =u i+1 / a i+1 ;

[0119] (3.3.6) Compare whether the gas Mach number reaches the nozzle Mach number. nozzle If the target is not reached, repeat steps (3.3.4)-(3.3.5) until the nozzle Mach number is reached. nozzle .

[0120] (3.3.7) Based on the pressure P of the last iteration step i+1 The static temperature T of the free flow field gas at the nozzle exit is calculated from the entropy value S. flow Static pressure P flow .

[0121] The calculation method for the oblique shock wave post-parameters in steps (1.2) and (3.4) above is as follows:

[0122] The relationship between the tangential and normal parameters before and after the oblique shock wave is shown in the figure below. Figure 2 As shown in the figure, we can see that:

[0123] w1 = u1sinβ;

[0124] w2 = u2sin(β-θ);

[0125] v1 = u1cosβ;

[0126] v2 = u2cos(β-θ);

[0127] Where β is the oblique shock wave angle, θ is the inlet compression angle, w and v are the normal and tangential velocity components of velocity u, respectively, subscript 1 indicates the parameters before the oblique shock wave, and subscript 2 indicates the parameters after the oblique shock wave.

[0128] Since v1 = v2, then:

[0129] w2 = u1cosβtan(β-θ);

[0130] From the above relationships, the shock angle and compression angle can be expressed as functions of the incoming tangential velocity and normal velocity, i.e.:

[0131] β=sin -1 (w1 / u1);

[0132]

[0133] The parameters before and after the shock wave simultaneously satisfy the following conservation relationships of mass, momentum, and energy:

[0134] ρ1w1=ρ2w2;

[0135]

[0136] In the formula, ρ, P, and H represent the gas density, pressure, and enthalpy, respectively.

[0137] Based on the above relationships between the tangential and normal parameters before and after the oblique shock wave, as well as the conservation relationships of mass, momentum, and energy, the calculation steps for the parameters after the oblique shock wave are as follows:

[0138] (1.2.01) Given the incoming flow parameters T1, P1, u1 and compression angle θ of the oblique shock wave;

[0139] (1.2.02) Given the initial value of the oblique shock wave angle β, calculate the normal velocity component w1 before the oblique shock wave;

[0140] (1.2.03) Given the initial value of the density ρ2 behind the oblique shock wave, calculate the specific volume, the temperature and pressure behind the wave:

[0141]

[0142] In the formula, V is the specific volume, P is the pressure, and ρ is the density;

[0143] (1.2.04) Calculate the normal velocity after the oblique shock wave and the difference function of momentum and energy before and after the oblique shock wave:

[0144]

[0145] In the formula, P and H are the momentum and energy difference functions before and after the oblique shock wave, respectively.

[0146] (1.2.05) Keeping the specific volume constant, calculate H(T2+ΔT) and P(T2+ΔT) under the condition of temperature disturbance ΔT;

[0147] (1.2.06) Keeping the temperature constant, calculate H(V2+ΔV) and P(V2+ΔV) under specific volume disturbance ΔV;

[0148] (1.2.07) Calculate the Jacobian matrix using first-order difference:

[0149]

[0150] (1.2.08) Calculate the system of linear equations Determine the temperature and specific volume corrections δT and δV.

[0151] (1.2.09) Appropriately restrict δT and δV: when |δT|>0.2*T², take 0.2*T²*sgn(δT); when |δV|>0.2*V², and The value is 0.2*V2*sgn(δV) when |δV|>0.5*(V1-V2) and when V2+δV>V1, the value is 0.5*(V1-V2)*sgn(δV), where sgn is the sign function.

[0152] (1.2.10) Determine the new oblique shock wave back temperature T2' and specific volume V2':

[0153] T′2=T2-δT;

[0154] V′2=V2-δV;

[0155] (1.2.11) Check if the convergence condition has been met:

[0156] T′2-T2<T Err ;

[0157] V′2-V2<V Err ;

[0158] In the formula, T Err V Err Allowable errors for the set temperature and specific volume.

[0159] (1.2.12) Repeat steps (1.2.03)-(1.2.11) until the convergence condition is met;

[0160] (1.2.13) Determine the pressure, velocity and Mach number of the oblique shock wave based on the temperature and specific volume behind the oblique shock wave.

[0161] The following are typical examples:

[0162] For a real engine with a 12.5° inlet on two compression surfaces and oblique detonation, at a flight altitude of 40km and a Mach number of 10, the conventional wind tunnel test scheme is as follows: Figure 3 As shown, the wind tunnel free flow field requires a Mach number of 10, and the wind tunnel experimental model must have the same number of inlet stages and compression angle as the real engine. Since large shock tunnels with a Mach number of Ma10 are rare, this patented method is used to calculate the equivalent wind tunnel experimental state to simulate this condition using a low Mach number wind tunnel. According to the above calculation method, the inlet velocity of the combustion chamber of the oblique detonation engine in flight mode is determined to be 2895 m / s, the pressure to be 13.37 kPa, and the temperature to be 1057 K. Conventional experimental methods require a Mach 10 wind tunnel. Using the above method, the calculated equivalent wind tunnel experimental state is a Mach 6 wind tunnel with a total temperature of 3800 K, a total pressure of 6.4 MPa, and a 15° single-compression-face inlet. The equivalent combustion chamber inlet velocity is 2631 m / s, the pressure to be 13.68 kPa, and the temperature to be 1024 K. The equivalent state differs from the actual state of the oblique detonation engine in the following ways: the velocity difference is less than 9.1%, the pressure difference is less than 2.3%, and the temperature difference is less than 3.1%, which meets the requirements of engineering experiments. This enables the simulation of high Mach number states in a low Mach number wind tunnel, expands the flow field conditions for conducting wind tunnel experiments on oblique detonation engines, and is conducive to promoting the engineering development of oblique detonation engines.

[0163] This invention can utilize a low Mach number wind tunnel to simulate the inlet parameters of the combustion chamber of a slant detonation engine under high Mach number conditions, thereby broadening the experimental conditions for conducting wind tunnel experiments on the combustion and thrust performance of slant detonation engines and promoting the engineering development of slant detonation engines.

[0164] Finally, it should be noted that the above embodiments are only used to illustrate the technical solutions of the present invention, and not to limit them; although the present invention has been described in detail with reference to the foregoing embodiments, those skilled in the art should understand that modifications can still be made to the technical solutions described in the foregoing embodiments, or equivalent substitutions can be made to some or all of the technical features; and these modifications or substitutions do not cause the essence of the corresponding technical solutions to deviate from the scope of the technical solutions of the embodiments of the present invention.

Claims

1. A wind tunnel experimental method for the equivalent state of a hypersonic oblique detonation engine, characterized in that, Includes the following steps: (1) Obtain the inlet parameters of the combustion chamber of the oblique detonation experimental model; the specific process is as follows: (1.1) Define the wind tunnel test mission for the oblique detonation engine and determine the flight Mach number of the oblique detonation test model according to the requirements of the test mission. Flight status, high-altitude static temperature static pressure and the compression angles at each stage of the intake manifold; (1.2) The Mach number at the combustion chamber inlet during flight was calculated using the method of calculating the aftershock parameters of the oblique shock wave. static temperature at the combustion chamber inlet during flight and the static pressure at the combustion chamber inlet during flight ; (2) Determine the flow field simulation capability of the experimental wind tunnel; specifically including the following steps: (2.1) Determine the Mach number of the wind tunnel nozzle The range of variation; (2.2) Determine the range of variation of total temperature T0 and total pressure P0 in the wind tunnel flow field; (3) Calculate the inlet parameters of the equivalent combustion chamber and determine the equivalent experimental scheme; the specific process is as follows: (3.1) Determine the allowable percentage differences in static temperature, static pressure, and Mach number between the equivalent state and the flight state, T%, P%, and Ma%, and give the compression angle settings for each stage of the equivalent inlet. (3.2) Given the total temperature, total pressure, and Mach number of the flow field; (3.3) Calculate the static temperature of the free flow field at the nozzle exit based on the isentropic expansion relationship. static pressure ; (3.4) Based on the given first-stage compression angle of the inlet duct and the calculated static temperature of the free flow field at the nozzle exit static pressure The flow field parameters after the first stage compression in the inlet were calculated using the oblique shock wave post-parameter calculation method. (3.5) Repeat step (3.4) to complete the parameter calculation after compression of each stage of the intake, and obtain the equivalent parameters of the combustion chamber inlet of the oblique detonation engine: equivalent state combustion chamber inlet static temperature Equivalent state combustion chamber inlet static pressure and equivalent state combustion chamber inlet Mach number ; (3.6) Calculate whether the differences in static temperature, static pressure, and Mach number at the combustion chamber inlet under equivalent and flight conditions satisfy the following three conditions: ； ； ； (3.7) If the three conditions in step (3.6) above are met, stop the calculation; otherwise, change the total temperature, total pressure and Mach number of the free flow field at the nozzle exit, and repeat steps (3.2)-(3.6) above to complete the calculation under all experimental wind tunnel flow field simulation capability conditions.

2. The wind tunnel experimental method for the equivalent state of a hypersonic oblique detonation engine as described in claim 1, characterized in that, The specific process of step (3.3) is as follows: (3.3.1) The wind tunnel accelerates stationary gas through isentropic expansion via nozzles to form a free flow field with a certain velocity at the nozzle exit. The stationary gas is in the stagnation chamber of the wind tunnel, and the entropy value remains unchanged during the isentropic expansion process. Given the total temperature P0, total pressure T0, and nozzle exit Mach number of the flow field. ; （ 3.3.2) Calculate the entropy S0 and enthalpy H0 of the resident gas; ； ； In the formula, R is the gas constant, and c1, c2, c3, c4, c5, b1, and b2 are piecewise constants with respect to temperature; (3.3.3) Set the pressure expansion rate dP during the isentropic expansion process; (3.3.4) The isentropic expansion process is calculated in multiple iterative steps, with the pressure change at each step being... Since the entropy S0 remains constant due to isentropic expansion, calculate the local gas enthalpy H. i+1 and speed of sound a i+1 : ； ； In the formula, Specific heat ratio of gases; Due to energy conservation during nozzle expansion, the local gas velocity u i+1 : ； (3.3.5) The local gas Mach number is: ； (3.3.6) Compare whether the gas Mach number reaches the nozzle Mach number. If the target is not reached, repeat steps (3.3.4)-(3.3.5) until the nozzle exit Mach number is reached. ; (3.3.7) Based on the pressure of the last iteration step The static temperature of the free-flowing gas at the nozzle exit was calculated using the entropy value S0. static pressure .

3. The wind tunnel experimental method for the equivalent state of a hypersonic oblique detonation engine as described in claim 1, characterized in that, The calculation method for the oblique shock wave post-parameters in steps (1.2) and (3.4) is as follows: The forward and backward tangential and normal parameters of the oblique shock wave satisfy the following relationship: ； ； ； ； Where β is the oblique shock wave angle, θ is the inlet compression angle, w and v are the normal and tangential velocity components of velocity u, respectively, subscript 1 indicates the parameters before the oblique shock wave, and subscript 2 indicates the parameters after the oblique shock wave; because ,but: ； From the above relationships, it can be seen that the shock wave angle and compression angle are functions of the incoming tangential velocity and normal velocity, respectively: ； ； The parameters before and after the shock wave simultaneously satisfy the following conservation relationships of mass, momentum, and energy: ； ； ； In the formula, ρ, P, and H represent the gas density, pressure, and enthalpy, respectively. Based on the above relationships between the tangential and normal parameters before and after the oblique shock wave, as well as the conservation relationships of mass, momentum, and energy, the parameters after the oblique shock wave are calculated. The specific process is as follows: (1.2.01) Given the incoming flow parameters T1, P1, u1 and compression angle θ of the oblique shock wave; (1.2.02) Given the initial value of the oblique shock wave angle β, calculate the normal velocity component w1 before the oblique shock wave; (1.2.03) Given the initial value of the back density of the oblique shock wave. Calculate the specific volume and the pressure behind the wave: ； ； In the formula, V is the specific volume, P is the pressure, and ρ is the density; (1.2.04) Calculate the normal velocity after the oblique shock wave and the difference function of momentum and energy before and after the oblique shock wave: ； ； ； In the formula, P and H are the momentum and energy difference functions before and after the oblique shock wave, respectively; (1.2.05) Keeping the specific volume constant, calculate the temperature disturbance ΔT. and ; (1.2.06) Keeping the temperature constant, calculate the specific volume perturbation. Below and ; (1.2.07) Calculate the Jacobian matrix using first-order difference: ； ； ； ； (1.2.08) Calculate the system of linear equations Determine the temperature and specific volume correction amount and ; (1.2.09) to and To impose restrictions, when At that time, Values ;when and At that time, Values ,when And satisfy At that time, Values In the formula, sgn is the sign function; (1.2.10) Determine the new oblique shock wave back temperature and specific volume : ； ； (1.2.11) Check if the convergence condition has been met: ； ； In the formula, T Err V Err Allowable errors for the set temperature and specific volume; (1.2.12) Repeat steps (1.2.03) to (1.2.11) until the convergence condition is met; (1.2.13) Determine the pressure, velocity and Mach number of the oblique shock wave based on the temperature and specific volume behind the oblique shock wave.

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