Hypersonic speed oblique detonation engine equivalent state wind tunnel experiment method

Abstract

The invention provides an equivalent state wind tunnel experiment method for a hypersonic speed oblique detonation engine. The method mainly comprises the following steps: (1) obtaining inlet parameters of a combustion chamber of an oblique detonation experiment model; (2) determining the simulation capability of the experimental wind tunnel flow field; and (3) calculating equivalent combustion chamber inlet parameters, and determining an equivalent experiment scheme. According to the invention, the low-Mach-number wind tunnel can be utilized to simulate the inlet parameters of the combustion chamber of the oblique detonation engine in the high-Mach-number state, so that the experiment conditions are widened for carrying out the combustion and thrust performance wind tunnel experiment of the oblique detonation engine, and the engineering development of the oblique detonation engine is promoted.

Description

Technical Field

[0001] The present invention relates to the technical field of oblique detonation engines, and particularly relates to a method for equivalent state wind tunnel experiments of a hypersonic oblique detonation engine. Background Art

[0002] The oblique detonation engine is a new type of power system for air-breathing hypersonic vehicles. It utilizes an oblique detonation wave to achieve efficient combustion in a supersonic airflow, and has technical advantages such as fast heat release, high specific impulse, short combustion chamber, and few moving parts. Thanks to the propagation speed of the detonation wave on the order of kilometers per second, the inlet airflow of the oblique detonation engine combustion chamber can maintain a high speed, and strong shock ignition can achieve the conversion of fuel chemical energy into internal energy within a millimeter scale, greatly shortening the combustion chamber size and having potential engineering advantages. It is a frontier research content of hypersonic propulsion technology. Theoretically, the thrust performance of an engine based on detonation combustion can be more than 30% higher than that of existing scramjet engines based on isobaric combustion. Through an effective combustion organization method, both entropy increase and total pressure loss can be further reduced, thereby further optimizing the thrust performance of the engine. It is considered a more suitable air-breathing propulsion method for hypersonic vehicles with a Mach number above 9, and can break through the Mach number limit of current hydrocarbon fuel scramjet engines. Since there are very complex physical and chemical processes in the oblique detonation engine combustion chamber, in-depth research on the oblique detonation combustion mechanism such as the initiation characteristics, unsteady characteristics, and wave surface stability of the oblique detonation wave is the key basis for the engineering development of oblique detonation engines. Currently, the numerical simulation research on the oblique detonation combustion phenomenon and mechanism has been relatively in-depth, but the engineering development of oblique detonation engines is still in its infancy, and it is urgent to carry out wind tunnel experiments to verify the combustion performance and component performance of oblique detonation engines to promote the integrated design of the entire oblique detonation engine. The main problems in conducting wind tunnel experiments on oblique detonation engines are as follows: The oblique detonation engine is applicable to flight conditions with a Mach number greater than 9. The conventional wind tunnel experiment method is that the Mach number of the wind tunnel flow field is the same as the flight Mach number of the engine, and the engine model is placed in the flow field for testing. The main difficulties faced in conducting wind tunnel experiment research on oblique detonation engines using this scheme are as follows: Since the oblique detonation chemical reaction process is closely related to the incoming flow temperature, pressure, and model scale ratio, in order to truly reflect the chemical reaction process of oblique detonation during flight, it is required that the total incoming flow temperature, total pressure of the wind tunnel experiment, and the scale of the wind tunnel experiment model are the same as those of the real aircraft, resulting in a high total incoming flow temperature and total pressure in the wind tunnel and a large scale of the wind tunnel experiment flow field; on the other hand, when using aviation kerosene as fuel, due to the increase in fuel ignition delay time, a longer effective experiment time is required for the wind tunnel. Currently, the number of wind tunnels with a Mach number above 9 that can meet the experimental conditions of oblique detonation engines is small, which is not conducive to the development and testing of oblique detonation engines. Summary of the Invention

[0003] In view of the technical problems existing in the above-mentioned background art, the present invention proposes a method for equivalent state wind tunnel experiments of a hypersonic oblique detonation engine, which can use a low Mach number wind tunnel to simulate the inlet parameters of the combustion chamber of an oblique detonation engine under high Mach number conditions, thereby broadening the experimental conditions for wind tunnel experiments on the combustion and thrust performance of oblique detonation engines and promoting the engineering development of oblique detonation engines.

[0004] To solve the above technical problems, a method for equivalent state wind tunnel experiments of a hypersonic oblique detonation engine provided by the present invention mainly includes the following steps:

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

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

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

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

[0009] (1.1) Define the wind tunnel experiment task of the oblique detonation engine, and determine the flight Mach number Ma fly 、flight altitude static temperature T fly 、static pressure P fly and the compression angles of each stage of the inlet duct according to the requirements of the experiment task;

[0010] (1.2) Calculate the flight state combustion chamber inlet Mach number Ma ode 、flight state combustion chamber inlet static temperature T ode and flight state combustion chamber inlet static pressure P ode by using the calculation method of the parameters behind the oblique shock wave.

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

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

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

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

[0015] (3.1) Determine the allowable difference percentage values T%, P% and Ma% of the static temperature, static pressure and Mach number between the equivalent state and the flight state, and set the compression angles of each stage of the equivalent inlet duct;

[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 according to the isentropic expansion relationship flow and static pressure P flow ;

[0018] (3.4) According to 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 and static pressure P flow , use the calculation method of the parameters after the oblique shock wave to calculate the flow field parameters after the first-stage compression of the inlet;

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

[0020] (3.6) Calculate whether the differences in the static temperature, static pressure, and Mach number at the inlet of the combustion chamber between the equivalent state and the flight state meet 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 the above steps (3.2)-(3.6) to complete the calculation under all experimental wind tunnel flow field simulation capabilities.

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

[0024] (3.3.1) The wind tunnel accelerates the static gas through isentropic expansion through the nozzle to form a free flow field with a certain velocity at the nozzle exit. The gas in the static state is in the stagnation chamber of the wind tunnel, and the entropy value remains unchanged during the isentropic expansion process; given the total pressure P0, total temperature T0 of the flow field, and Mach number Ma at the nozzle exit nozzle ;

[0025] (3.3.2) Calculate the entropy value S0 and enthalpy value H0 of the gas in the stagnation chamber;

[0026]

[0027] In the formula, R is the gas constant, and c1, c2, c3, c4, c5, b1, 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) Calculate the isentropic expansion process through multi-step iteration, with the pressure change per step being P i+1 = P i - dP. Due to isentropic expansion, the entropy value S remains unchanged. Calculate the local gas enthalpy value H i+1 and the speed of sound a i+1 :

[0030]

[0031] where γ is the specific heat ratio of the gas;

[0032] Due to energy conservation during the nozzle expansion process, 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 Ma nozzle . If not, repeat steps (3.3.4)-(3.3.5) for calculation until the nozzle exit Mach number Ma nozzle is reached;

[0037] (3.3.7) Calculate the static temperature T i+1 and static pressure P flow of the free stream gas at the nozzle exit based on the pressure P flow in the last iteration step and the entropy value S.

[0038] For the experimental method of the hypersonic oblique detonation engine equivalent state wind tunnel, wherein the calculation method of the parameters after the oblique shock wave in steps (1.2) and (3.4) is:

[0039] The tangential and normal parameters before and after the oblique shock wave satisfy the following relationships:

[0040] w1 = u1sinβ;

[0041] w2 = u2sin(β - θ);

[0042] v1 = u1cosβ;

[0043] v2 = u2cos(β - θ);

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

[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 flow tangential velocity and normal velocity, that is:

[0048] β = sin -1 (w1 / u1);

[0049]

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

[0051] ρ1w1 = ρ2w2;

[0052]

[0053] Where ρ, P, and H are the gas density, pressure, and enthalpy value respectively;

[0054] According to the relationships between the tangential and normal parameters before and after the oblique shock and the mass, momentum, and energy conservation relationships, calculate the parameters after the oblique shock. The specific process is as follows:

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

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

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

[0058]

[0059] Where 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 and the momentum and energy difference functions before and after the oblique shock:

[0061]

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

[0063] (1.2.05) Keep the specific volume unchanged, and calculate H(T2 + ΔT) and P(T2 + ΔT) under the condition of temperature perturbation ΔT;

[0064] (1.2.06) Keep the temperature unchanged, and calculate Η(V2 + ΔV) and P(V2 + ΔV) under the specific volume perturbation ΔV;

[0065] (1.2.07) Use the first-order difference to calculate the Jacobian matrix:

[0066]

[0067] (1.2.08) Calculate the linear equations Determine the temperature and specific volume correction amounts δT and δV;

[0068] (1.2.09) Impose restrictions on δT and δV. When |δT| > 0.2 * T₂, take 0.2 * T₂ * sgn(δT); when |δV| > 0.2 * V₂ and take the value of 0.2 * V₂ * sgn(δV) at this time, and when |δV| > 0.5 * (V₁ - V₂) and V₂ + δV > V₁, take the value of 0.5 * (V₁ - V₂) * sgn(δV), where sgn is the sign function;

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

[0070] T′₂ = T₂ - δT;

[0071] V′₂ = V₂ - δV;

[0072] (1.2.11) Check whether the convergence condition is reached:

[0073] T′₂ - T₂ < T Err ;

[0074] V′₂ - V₂ < V Err ;

[0075] In the formula, T Err 、V Err are the set allowable errors of temperature and specific volume;

[0076] (1.2.12) Repeat the above steps (1.2.03) - (1.2.11) until the convergence condition is satisfied;

[0077] (1.2.13) Determine the oblique shock wave pressure, velocity and Mach number according to the post-oblique shock wave temperature and specific volume.

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

[0079] The equivalent state wind tunnel experiment method of the hypersonic oblique detonation engine of the present invention can use a low Mach number wind tunnel to simulate the inlet parameters of the combustion chamber of the oblique detonation engine under high Mach number conditions, so as to broaden the experimental conditions for carrying out wind tunnel experiments on the combustion and thrust performance of the oblique detonation engine. Compared with the conventional method, it can reduce the requirement for the wind tunnel driving ability, expand the working condition test range of the oblique detonation engine, and promote the engineering development of the oblique detonation engine. BRIEF DESCRIPTION OF THE DRAWINGS

[0080] In order to more clearly illustrate the specific embodiments of the present invention or the technical solutions in the prior art, the following will briefly introduce the drawings required for the description of the specific embodiments or the prior art. Obviously, the drawings in the following description are some embodiments of the present invention. For those of ordinary skill in the art, without creative efforts, other drawings can also be obtained based on these drawings.

[0081] Figure 1 is a flow chart of the equivalent state wind tunnel experiment method of the hypersonic oblique detonation engine of the present invention;

[0082] Figure 2 is a diagram showing the relationship between the tangential and normal parameters before and after the oblique shock wave involved in the equivalent state wind tunnel experiment method of the hypersonic oblique detonation engine of the present invention;

[0083] Figure 3 is a conventional wind tunnel experiment scheme for a two-stage compression inlet engine at Mach 10;

[0084] Figure 4 is an equivalent wind tunnel experiment scheme for conducting experiments at Mach 6 with a single-stage compression inlet. DETAILED DESCRIPTION OF THE EMBODIMENTS

[0085] The following will clearly and completely describe the technical solutions of the present invention with reference to the accompanying drawings. Obviously, the described embodiments are some, but not all, of the embodiments of the present invention. All other embodiments obtained by those of ordinary skill in the art without creative efforts based on the embodiments of the present invention belong to the scope of protection of the present invention.

[0086] The following further explains and illustrates the present invention in conjunction with specific embodiments.

[0087] The conventional experimental method of hypersonic oblique detonation engine in wind tunnel is that the Mach number of the wind tunnel flow field is the same as that of the engine in flight, and the engine model is placed in the flow field for testing. From the working principles of the wind tunnel and the oblique detonation engine, it can be known that: the wind tunnel flow field is formed by the isentropic expansion of the gas in the wind tunnel settling chamber through the nozzle; the oblique detonation engine mainly consists of an inlet, a combustion chamber and a tail nozzle. The main function of the inlet is to compress the incoming air to increase its pressure and temperature, creating conditions for the oblique detonation combustion in the combustion chamber. Therefore, in the conventional oblique detonation wind tunnel experiment, the air first expands through the nozzle and then is compressed by the oblique detonation engine inlet to form the incoming flow conditions required by the combustion chamber. Since expansion and compression are opposite physical processes, by reducing the expansion process of the wind tunnel nozzle and optimizing the design of the inlet compression angle, it is possible to simulate the inlet parameters of the combustion chamber of the oblique detonation engine in a high Mach number state in a low Mach number wind tunnel, thus expanding the experimental conditions for the combustion and thrust performance wind tunnel experiments of the oblique detonation engine.

[0088] As Figure 1 shown, a hypersonic oblique detonation engine equivalent state wind tunnel experimental method provided by 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 experimental tasks of the oblique detonation engine, and determine the flight Mach number Ma fly , the high-altitude static temperature T fly , the static pressure P fly of the flight state, and the compression angles of each stage of the inlet;

[0091] (1.2) Use the calculation method of the parameters behind the oblique shock wave to calculate the Mach number Ma ode , the static temperature T ode at the inlet of the combustion chamber in the flight state, and the static pressure P ode at the inlet of the combustion chamber in the flight state;

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

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

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

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

[0096] (3.1) Determine the allowable difference percentage values T%, P% and Ma% of the static temperature, static pressure and Mach number between the equivalent state and the flight state, and set the compression angles of 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 according to the isentropic expansion relationship flow and static pressure P flow

[0099] (3.4) According to 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 and static pressure P flow , calculate the flow field parameters after the first-stage compression of the inlet using the calculation method for the parameters after an oblique shock wave;

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

[0101] (3.6) Calculate whether the differences in the static temperature, static pressure, and Mach number at the inlet of the combustion chamber between the equivalent state and the flight state 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 the above step (3.3) is as follows:

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

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

[0107] (3.3.2) Calculate the entropy value S0 and enthalpy value H0 of the gas in the stagnation chamber;

[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) Calculate the isentropic expansion process through multi-step iteration, with the pressure change per step being P i+1 = P i - dP. Due to isentropic expansion, the entropy value S remains unchanged. Calculate the local gas enthalpy value H i+1 and the speed of sound a i+1 ,

[0112]

[0113]

[0114] where γ is the specific heat ratio of the gas.

[0115] Due to energy conservation during the nozzle expansion process, 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 Ma nozzle . If not, repeat steps (3.3.4) - (3.3.5) for calculation until the nozzle Mach number Ma nozzle is reached.

[0120] (3.3.7) Calculate the static temperature T i+1 and static pressure P flow of the free flow field gas at the nozzle exit based on the pressure P flow and entropy value S of the last iteration step.

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

[0122] The relationship diagram of the tangential and normal parameters before and after the oblique shock wave is as Figure 2 shown. It can be seen from the figure that:

[0123] w1 = u1sinβ;

[0124] w2 = u2sin(β - θ);

[0125] v1 = u1cosβ;

[0126] v2 = u2cos(β - θ);

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

[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 flow tangential velocity and normal velocity, i.e.:

[0131] β = sin -1 (w1 / u1);

[0132]

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

[0134] ρ1w1 = ρ2w2;

[0135]

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

[0137] According to the above relationships between the tangential and normal parameters before and after the oblique shock and the mass, momentum, and energy conservation relationships, the calculation steps for calculating the parameters after the oblique shock are as follows:

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

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

[0140] (1.2.03) Given the initial value of the density ρ2 after the oblique shock wave, calculate the specific volume and the temperature and pressure after 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 and the momentum and energy difference functions before and after the oblique shock:

[0144]

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

[0146] (1.2.05) Keep the specific volume unchanged, and calculate H(T2 + ΔT) and P(T2 + ΔT) under the condition of temperature perturbation ΔT;

[0147] (1.2.06) Keep the temperature unchanged, and calculate Η(V2 + ΔV) and P(V2 + ΔV) under the specific volume perturbation ΔV;

[0148] (1.2.07) Use the first-order difference to calculate the Jacobian matrix:

[0149]

[0150] (1.2.08) Calculate the linear equations Determine the temperature and specific volume correction amounts δT and δV.

[0151] (1.2.09) Apply appropriate restrictions to δT and δV. When |δT| > 0.2 * T2, take 0.2 * T2 * sgn(δT); when |δV| > 0.2 * V2 and take the value of 0.2 * V2 * sgn(δV) at this time, and when |δV| > 0.5 * (V1 - V2) and V2 + δV > V1, take the value of 0.5 * (V1 - V2) * sgn(δV), where sgn is the sign function.

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

[0153] T′2 = T2 - δT;

[0154] V′2 = V2 - δV;

[0155] (1.2.11) Check whether the convergence condition is reached:

[0156] T′2 - T2 < T Err ;

[0157] V′2 - V2 < V Err ;

[0158] Wherein, T Err , V Err are the set allowable errors of temperature and specific volume.

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

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

[0161] The following are typical examples:

[0162] For the real engine of the 12.5° inlet oblique detonation with two compression surfaces, the flight altitude is 40 km and the flight Mach number Ma is 10. The conventional wind tunnel experiment scheme is as Figure 3 shown. It is required that the Mach number of the wind tunnel free stream field is 10, and the wind tunnel experiment model has the same number of inlet stages and compression angle as the real engine. Since there are very few large shock wind tunnels with a Mach number Ma of 10, the method of this patent is used to calculate the equivalent wind tunnel experiment state for simulating this condition using a low Mach number wind tunnel. According to the above calculation method, the flow field velocity at the combustion chamber inlet of the oblique detonation engine in the flight state is determined to be 2895 m / s, the pressure is 13.37 kPa, and the temperature is 1057 K. The conventional experimental method needs to be carried out in a Ma10 wind tunnel. By using the above method for calculation, the result of calculating the equivalent wind tunnel experiment state is a wind tunnel with a Mach number of 6, a total temperature of 3800 K and a total pressure of 6.4 Mpa in the stagnation chamber, and a 15° single compression surface inlet. The inlet velocity of the combustion chamber in the equivalent state is 2631 m / s, the pressure is 13.68 kPa, and the temperature is 1024 K. The differences between the equivalent state and the real state of the oblique detonation engine are: 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%, meeting the requirements of engineering experiments. Thus, the simulation of the high Mach number state by a low Mach number wind tunnel is realized, the flow field conditions for carrying out the wind tunnel experiment of the oblique detonation engine are expanded, and it is beneficial to promote the engineering development of the oblique detonation engine.

[0163] The present invention can use a low Mach number wind tunnel to simulate the inlet parameters of the combustion chamber of an oblique detonation engine in a high Mach number state, thereby broadening the experimental conditions for carrying out the wind tunnel experiments on the combustion and thrust performance of the oblique detonation engine and promoting the engineering development of the oblique detonation engine.

[0164] Finally, it should be noted that: the above embodiments are only used to illustrate the technical solutions of the present invention, rather than to limit them; although the present invention has been described in detail with reference to the foregoing embodiments, those of ordinary skill in the art should understand that: they can still modify the technical solutions recorded in the foregoing embodiments, or perform equivalent replacements on some or all of the technical features; and these modifications or replacements do not make the essence of the corresponding technical solutions deviate from the scope of the technical solutions of the embodiments of the present invention.

Claims

1. An experimental method for equivalent state wind tunnel of hypersonic oblique detonation engine, characterized in that, It mainly includes the following steps: (1) Obtain the inlet parameters of the combustion chamber of the oblique detonation experiment model; (2) Determine the simulation ability of the experimental wind tunnel flow field; (3) Calculate the equivalent combustion chamber inlet parameters and determine the equivalent experimental scheme.

2. The equivalent state wind tunnel experimental method for a hypersonic oblique detonation engine according to claim 1, characterized in that The specific process of step (1) is as follows: (1.1) Define the wind tunnel experiment tasks of the oblique detonation engine, and determine the flight Mach number Ma of the oblique detonation experiment model according to the requirements of the experiment tasks fly , the high-altitude static temperature T of the flight state fly , the static pressure P fly and the compression angles of each stage of the inlet; (1.2) Calculate the Mach number Ma at the combustor inlet in the flight state using the calculation method for the parameters behind the oblique shock wave ode , the static temperature T at the combustor inlet in the flight state ode and the static pressure P at the combustor inlet in the flight state ode .

3. The hypersonic oblique detonation engine equivalent state wind tunnel experiment method according to claim 1, characterized in that Step (2) specifically includes the following steps: (2.1) Determine the variation range of the Mach number Ma of the wind tunnel nozzle nozzle ; (2.2) Determine the variation ranges of the total temperature T0 and total pressure P0 of the wind tunnel flow field.

4. The experimental method of the hypersonic oblique detonation engine equivalent state wind tunnel according to claim 1, characterized in that, The specific process of step (3) is as follows: (3.1) Determine the allowable difference percentage values T%, P%, and Ma% of the static temperature, static pressure, and Mach number between the equivalent state and the flight state, and set the compression angles of each stage of the equivalent inlet; (3.2) Specify the total temperature, total pressure, and Mach number of the flow field; (3.3) Calculate the static temperature T of the free flow field at the nozzle outlet according to the isentropic expansion relationship flow and the static pressure P flow ; (3.4) According to 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 and static pressure P flow , use the calculation method of the parameters after the oblique shock wave to calculate the flow field parameters after the first-stage compression of the inlet; (3.5) Repeat step (3.4) to complete the parameter calculation after the compression of each stage of the inlet, and obtain the equivalent parameters at the inlet of the scramjet engine combustion chamber: the static temperature T at the inlet of the equivalent state combustion chamber equi , the static pressure P at the inlet of the equivalent state combustion chamber equi and the Mach number Ma at the inlet of the equivalent state combustion chamber equi ; (3.6) Calculate whether the differences in the static temperature, static pressure, and Mach number at the combustion chamber inlet between the equivalent state and the flight state meet the following three conditions: (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 outlet, and repeat steps (3.2)-(3.6) above to complete the calculation under all experimental wind tunnel flow field simulation ability conditions.

5. The experimental method of the equivalent state wind tunnel for the hypersonic oblique detonation engine according to claim 4, characterized in that, The specific process of step (3.3) is as follows: (3.3.1) The wind tunnel accelerates the static gas through isentropic expansion through the nozzle to form a free flow field at the nozzle outlet with a certain velocity. The gas in the static state 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 of the flow field and the Mach number Ma at the nozzle exit nozzle ; (3.3.2) Calculate the entropy value S0 and enthalpy value H0 of the stagnation chamber 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) Calculate the isentropic expansion process through multi-step iteration, with the pressure change per step being P i+1 = P i -dP. Due to isentropic expansion, the entropy value S remains unchanged. Calculate the local gas enthalpy value H i+1 and the speed of sound a i+1 : In the formula, γ is the specific heat ratio of the gas; Due to the conservation of energy during the expansion process of the nozzle, the local gas velocity u i+1 : u i+1 = sqrt(H i +(u i ) 2 / 2.0 - H i+1 ); (3.3.5) The local gas Mach number is: Ma i+1 = u i+1 / a i+1 ; (3.3.6) Compare whether the gas Mach number reaches the nozzle Mach number Ma nozzle If it does not reach, repeat steps (3.3.4)-(3.3.5) for calculation until the Mach number Ma at the nozzle outlet is reached nozzle ; (3.3.7) Calculate the static temperature T of the gas in the free flow field at the nozzle outlet based on the pressure P and entropy value S at the last iteration step i+1 and static pressure P flow of the gas in the free flow field at the nozzle outlet. flow .

6. The hypersonic oblique detonation engine equivalent state wind tunnel experiment method according to claim 1, characterized in that The calculation methods of the parameters after the oblique shock wave in step (1.2) and step (3.4) are as follows: The tangential and normal parameters before and after the oblique shock wave satisfy the following relationships: w1 = u1sinβ; w2 = u2sin(β - θ); v1 = u1cosβ; v2 = u2cos(β - θ); 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 the velocity u respectively, the subscript 1 represents the parameters before the oblique shock wave, and the subscript 2 represents the parameters after the oblique shock wave; Since v1 = v2, then: w2 = u1cosβtan(β - θ); From the above relationships, the shock wave angle and compression angle can be expressed as functions of the incoming flow tangential velocity and normal velocity, that is: β = sin -1 (w1 / u1); The parameters before and after the shock wave simultaneously satisfy the following mass, momentum, and energy conservation relationships: ρ1w1 = ρ2w2; In the formula, ρ, P, and H are the gas density, pressure, and enthalpy value respectively; According to the above relationships between the tangential and normal parameters before and after the oblique shock wave and the mass, momentum, and energy conservation relationships, calculate the parameters after the oblique shock wave. The specific process is as follows: (1.2.01) Specify the incoming flow parameters T1, P1, u1 before the oblique shock wave and the compression angle θ; (1.2.02) Specify the initial value of the oblique shock wave angle β, and calculate the normal velocity component w1 before the oblique shock wave; (1.2.03) Specify the initial value of the density ρ2 after the oblique shock wave, and calculate the specific volume and the pressure after the wave: Wherein, 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 momentum and energy difference functions before and after the oblique shock wave: Wherein, 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 unchanged, calculate H(T2 + ΔT) and P(T2 + ΔT) under the condition of temperature perturbation ΔT; (1.2.06) Keeping the temperature unchanged, calculate Η(V2 + ΔV) and P(V2 + ΔV) under the specific volume perturbation ΔV; (1.2.07) Use the first-order difference to calculate the Jacobian matrix: (1.2.08) Calculate the linear equations Determine the temperature and specific volume correction amounts δT and δV; (1.2.09) Restrict δT and δV. When |δT| > 0.2 * T2, take 0.2 * T2 * sgn(δT); when |δV| > 0.2 * V2 and take the value of 0.2 * V2 * sgn(δV), and when |δV| > 0.5 * (V1 - V2) and V2 + δV > V1, take the value of 0.5 * (V1 - V2) * sgn(δV), where sgn is the sign function; (1.2.10) Determine the new temperature T’2 and specific volume V’2 behind the oblique shock wave: T’2 = T2 - δT; V’2 = V2 - δV; (1.2.11) Check whether the convergence condition is reached: T’2 - T2 < T Err ; V’2 - V2 < V Err ; where T Err and V Err are the set allowable errors of temperature and specific volume; (1.2.12) Repeat the above steps (1.2.03) - (1.2.11) until the convergence condition is satisfied; (1.2.13) Determine the oblique shock wave pressure, velocity, and Mach number according to the temperature and specific volume behind the oblique shock wave.

Images