MODELING METHOD FOR INTEGRATED INTAKE/EXHAUST/ENGINE AERO PROPULSION SYSTEM WITH MULTIPLE GEOMETRIC PARAMETERS ADJUSTABLE

Abstract

A modeling method for an integrated intake/exhaust/engine aero propulsion system with multiple geometric parameters adjustable includes the following steps: establishing an inlet and nozzle model by quasi one-dimensional aerodynamic thermodynamics and the method for solving the excitation system on the basis of a traditional engine component-level model; adding an inlet and engine flow balance equation and an engine and nozzle flow balance equation to the engine model, and establishing a propulsion system model based on the iteration method; and integrating the design of geometric parameters of an inlet and a nozzle into the model to realize the design of structure sizes of an intake/exhaust system and the simultaneous adjustment of multiple parameters.

Claims

1. A modeling method for an integrated intake/exhaust/engine aero propulsion system with multiple geometric parameters adjustable, comprising the following steps: first, establishing an inlet and nozzle model by quasi one-dimensional aerodynamic thermodynamics and the method for solving the excitation system in further consideration of the influence of the shock structure and the drag of the inlet on the engine performance as well as the changing rule of the flow coefficient and the thrust coefficient of the nozzle under different working conditions on the basis of a traditional engine component-level model; then, adding an inlet and engine flow balance equation and an engine and nozzle flow balance equation to the engine model, and establishing a propulsion system model based on the iteration method; and finally, integrating the design of geometric parameters of an inlet and a nozzle into the engine model to realize the design of structure sizes of an intake/exhaust system and the simultaneous adjustment of multiple parameters; the specific steps are as follows: S1: building of quasi one-dimensional aerodynamic thermodynamic model in intake/exhaust system S1.1: according to the actual engine structure, determining the basic types of an inlet and a nozzle; S1.2: determining the structure parameters and the design operating points of the inlet, and establishing the corresponding relationship between the structure parameters of the inlet and the design parameters of the actual engine critical state through the two-dimensional plane geometry relationship; and determining the structure size parameters of a convergent-divergent nozzle based on the actual engine structure; S1.3: determining a designed shock system structure, and assuming that the inlet conditions are known, solving the total pressure recovery coefficient and the flow coefficient of the inlet under different inlet conditions by the method for solving the excitation system; and when the wavefront Mach number Ma.sub.f, the adiabatic exponent of gas k and the ramp angle δ are known, solving the shock wave angle β by iteration according to formula (1), and determining the total pressure loss coefficient σ and the wave rear Mach number Ma.sub.b of the shock wave according to formula (2) and formula (3): $\begin{array}{cc}\tan \delta =\frac{{\mathrm{Ma}}_f^2{\sin }^2\beta -1}{\left[{\mathrm{Ma}}_f^2\left(\frac{k+1}{2}-{\sin }^2\beta \right)+1\right]\tan \beta }& \left(1\right)\end{array}$ $\begin{array}{cc}\sigma =\frac{{\left[\frac{\left(k+1\right){\mathrm{Ma}}_f^2{\sin }^2\beta }{2+\left(k-1\right){\mathrm{Ma}}_f^2{\sin }^2\beta }\right]}^{\frac{k}{k-1}}}{{\left[\frac{2k}{k+1}{\mathrm{Ma}}_f^2{\sin }^2\beta \frac{k-1}{k+1}\right]}^{\frac{1}{k-1}}}& \left(2\right)\end{array}$ $\begin{array}{cc}{\mathrm{Ma}}_b^2=\frac{{\mathrm{Ma}}_f^2+\frac{2}{k-1}}{\frac{2k}{k-1}{\mathrm{Ma}}_f^2{\sin }^2\beta -1}+\frac{{\mathrm{Ma}}_f^2{\cos }^2\beta }{\frac{k-1}{2}{\mathrm{Ma}}_f^2{\sin }^2\beta +1}& \left(3\right)\end{array}$ S1.4: establishing the calculation formula of the subsonic drag of the engine model; the drag D.sub.add under the subsonic condition is mainly composed of additional drag, calculated through the loss of momentum of the airflow before the inlet lip in the horizontal direction, and expressed by formula (4): wherein T.sub.th, Ma.sub.th, A.sub.th and W.sub.a, th represent the throat temperature, the throat Mach number, the throat area and the throat flow, δ.sub.0 represents the total turning angle of the inlet, Ma.sub.0 represents the inlet Mach number of the inlet, A.sub.0 represents the inlet free flow tube area, and k represents the adiabatic exponent of gas; $\begin{array}{cc}{D}_{\mathrm{add}}=\frac{W_{a,\mathrm{th}}}{{\mathrm{kMa}}_0}\left[\frac{{\mathrm{Ma}}_0}{{\mathrm{Ma}}_{\mathrm{th}}}\sqrt{\frac{T_{\mathrm{th}}}{T_0}}\left(1+{\mathrm{kMa}}_{th}^2\right).Math.\cos {\delta }_0-\left(\frac{A_{\mathrm{th}}}{A_0}.Math.\cos {\delta }_0+{\mathrm{kMa}}_0^2\right)\right]& \left(4\right)\end{array}$ S1.5: establishing the calculation formula of the supersonic drag of the engine model; under the supersonic condition, the external drag of the inlet comprises additional drag and overflow drag; when the flow coefficient of the inlet is greater than or equal to the maximum flow coefficient, the operation is under the critical or supercritical condition, and the overflow drag is 0; when the flow coefficient of the inlet is less than the maximum flow coefficient, the operation is under the subcritical condition, the shock wave does not seal the inlet, and the overflow drag appears; and the calculation formula of the supersonic drag D.sub.add is expressed by formula (5), wherein H.sub.e1, H.sub.e2 and H.sub.e3 respectively represent vertical section heights of drag between shock waves of the inlet, P.sub.s1, P.sub.s2 and P.sub.s3 represent static pressures after shock waves, and P.sub.s0 represents the inlet total pressure of the inlet;
D.sub.add=(P.sub.s1−P.sub.s0)H.sub.e1+(P.sub.s2−P.sub.s0)H.sub.e2+(P.sub.s3−P.sub.s0)H.sub.e3  (5) S1.6: determining the basic type and adjustable variables of the nozzle, calculating the critical expansion ratio of the nozzle through structure parameters, and judging the operating state of the nozzle according to the total turbine outlet pressure and the environmental pressure: subcritical, critical and supercritical; and calculating the critical expansion ratio π.sub.NZ,cr of the nozzle according to formula (6), where Δ.sub.μk represents the flow coefficient component of the conical nozzle, which is related to the convergent half angle α and the length L.sub.c of the convergent section of the nozzle, and β is the divergent half angle; $\begin{array}{cc}{\pi }_{\mathrm{NZ},\mathrm{cr}}=1+\frac{{\left(\frac{k+1}{2}\right)}^{\frac{k}{k-1}}-1+29{\Delta }_{\mu k}}{1+0.088\frac{\sqrt{\stackrel{\_}{A_9}-1}}{0.005+{\beta }^{1.5}}},\stackrel{\_}{A_9}=\frac{A_9}{A_8}& \left(6\right)\end{array}$ S1.7: when the convergent-divergent nozzle is in the supercritical state, the area ratio of $\frac{A_9}{A_8}$ has an impact on the exit Mach number, wherein A.sub.9 represents the exit area of the nozzle, and A.sub.8 represents the throat area of the nozzle, obtaining the exit Mach number Ma.sub.9t by iterative solution according to formula (7); $\begin{array}{cc}\left(\frac{A_9}{A_8}\right)={\frac{1}{{\mathrm{Ma}}_{9t}}\left[\left(\frac{2}{\kappa +1}\right)\left(1+\frac{\kappa -1}{2}{\mathrm{Ma}}_{9t}^2\right)\right]}^{\frac{\left(\kappa +1\right)}{\left[2\left(\kappa -1\right)\right]}}& \left(7\right)\end{array}$ S1.8: calculating three characteristic flow state points of the convergent-divergent nozzle, determining the flow state in the nozzle according to the back pressure condition, and then calculating the exit total pressure, the static pressure, the total temperature and the flow rate of the nozzle; S1.9: calculating the flow coefficient Φ.sub.N and the thrust coefficient C.sub.F of the convergent-divergent nozzle according to the known parameters by means of an engineering empirical formula, which are used for calculating the actual throat flow and the actual thrust; formula (8) is the calculation method of the flow coefficient, wherein A.sub.7 represents the inlet area of the nozzle, and α represents the convergent half angle of the nozzle; and formula (9) is the calculation method of the thrust coefficient, wherein J.sub.c represents the impulse coefficient, J.sub.P(λ.sub.9) represents the computed impulse of the nozzle, and F.sub.N,id(π.sub.N,us) represents the ideal thrust of the nozzle; $\begin{array}{cc}{\Phi }_N=1-0.0585\frac{\left(1+2.63\alpha \right)\alpha }{1+{\alpha }^2}\left[1-{\left(\frac{A_8}{A_7}\right)}^2\right]-0.01\left[1-e^{\left(-0.5{\alpha }^2\right)}\right]\frac{A_8}{A_7}& \left(8\right)\end{array}$ $\begin{array}{cc}{C}_F=\frac{{\Phi }_N{\pi }_{N,\mathrm{us}}{J}_C{J}_P\left({\lambda }_9\right)-\frac{A_9}{A_7}}{{\Phi }_N{F}_{N,\mathrm{id}}\left({\pi }_{N,\mathrm{us}}\right)}& \left(9\right)\end{array}$ S2: establishment of component-level model of propulsion system S2.1: acquiring the characteristic curve of critical components of the aero-engine model; and respectively establishing the input/output module of a single component according to the sequence of propulsion system components based on aerodynamic thermodynamics, comprising gas flow equations and heat equations; S2.2: determining known input parameters of the model based on operating conditions and states of the model, determining the number and types of iteration variables through the common working equations, and conducting simulation calculation according to a gas process; S3: design of variable geometric parameters of inlet and nozzle S3.1: connecting the structure sizes of the inlet as input fixed parameters to an input end, wherein the values are determined by the design sizes; S3.2: connecting the rank angle, the bleed valve opening and the boundary layer suction opening of the inlet as variable parameters to the input end, wherein the parameters are adjusted at any time in a dynamic process; S3.3: connecting the inlet area, the length of the convergent section, the length of the divergent section, the convergent angle and the divergent angle of the nozzle as input fixed parameters to the input end; S3.4: connecting the throat area and the exit area of the nozzle as variable parameters to the input end; S4: building of integrated intake/exhaust/engine computing platform of supersonic vehicle S4.1: designing the inlet/exhaust/engine coupling component-level modeling and the iterative algorithm of supersonic vehicles by C++ programming, encapsulating the model through a dynamic link library, and introducing into a simulink module to establish a simulation platform; S4.2: the parameters of the input end of the platform comprise structure sizes and adjustable parameters of the inlet and the nozzle, adjustable parameters of the engine model and environmental operating conditions, establishing a simulation platform of a dynamic process.

Description

DESCRIPTION OF DRAWINGS

[0033] FIG. 1 is a schematic diagram of structure size parameters of a typical external compression inlet in a critical state.

[0034] FIG. 2 is a schematic diagram of structure size parameters of a typical convergent-divergent nozzle;

[0035] FIG. 3 is a flow chart of a characteristics calculation module of an inlet.

[0036] FIG. 4 is a schematic diagram of external drag calculation parameters of an inlet.

[0037] FIG. 5 is a flow chart of a calculation module of a convergent-divergent nozzle.

[0038] FIG. 6 is a flow chart of a component-level model of a typical propulsion system.

[0039] FIG. 7 shows the changing rule of the thrust performance of a propulsion system with a secondary rank angle δ.sub.2.

[0040] FIG. 8 shows the changing rule of the thrust performance of a propulsion system with a throat area A.sub.8.

[0041] FIG. 9 shows the changing rule of the thrust performance of a propulsion system with an exit area A.sub.9.

DETAILED DESCRIPTION

[0042] The embodiments of the present invention will be further described below in combination with the drawings and the technical solution.

[0043] S1: building of quasi one-dimensional aerodynamic thermodynamic model in intake/exhaust system

[0044] According to the actual engine structure, determining the types of an inlet and a nozzle, and determining design structure parameters of the inlet based on the critical operating state;

[0045] S1.1: determining the basic types of the inlet and the nozzle; in the embodiment, with a typical supersonic vehicle as an example, the inlet is an external compression inlet, and the nozzle is a convergent-divergent nozzle.

[0046] S1.2: determining the design operating points of the inlet. In the embodiment, an external compression inlet with two oblique and one direct shock waves is adopted, and the structure size parameters are determined through the two-dimensional plane geometry relationship to enable the shock waves to seal the inlet. The state is called a critical state, critical shock wave angles (β.sub.1des and β.sub.des) are determined by the structure size parameters, and the specific structure size parameters of the inlet are shown in FIG. 1. Assuming that the width of the inlet is S, and the lengths L.sub.1 and L.sub.2 and the height H.sub.c are size parameters, the capture area is A.sub.c=H.sub.c S;

[0047] S1.3: based on the actual engine structure, determining the structure size parameters (inlet area, length of convergent section, length of divergent section, convergent angle and divergent angle) of the convergent-divergent nozzle, and determining the adjustable range of adjustable parameters including throat area A.sub.8 and exit area A.sub.9; FIG. 2 is a structural schematic diagram of a convergent-divergent nozzle.

[0048] The present invention builds an inlet model based on the quasi one-dimensional calculation method, the basic calculation process of the inlet model is shown in FIG. 3, and the calculation thought is as follows:

[0049] S1.4: when the wavefront Mach number Ma.sub.f, the adiabatic exponent of gas k and the ramp angle δ are known, solving the shock wave angle β by iteration according to formula 1, and determining the total pressure loss coefficient σ and the wave rear Mach number Ma.sub.b of the shock wave according to formula 2 and formula 3. In the typical external compression inlet with two oblique and one direct shock waves, incoming flow passes through two oblique shock waves and one direct shock wave successively, and the above formulas are calculated for three times in sequence to obtain the shock wave angles β.sub.1 and β.sub.2 of the two oblique shock waves, the total pressure loss coefficients σ.sub.1, σ.sub.2 and σ.sub.3 of the three shock waves, and Mach number Ma.sub.3 after the direct shock wave; and the total pressure loss coefficient σ.sub.inlet of the inlet can be calculated according to formula 4 based on the above calculation results, and σ.sub.F represents the total pressure loss of wall friction;

[00007]$\begin{array}{cc}\tan \delta =\frac{{\mathrm{Ma}}_f^2{\sin }^2\beta -1}{\left[{\mathrm{Ma}}_f^2\left(\frac{k+1}{2}-{\sin }^2\beta \right)+1\right]\tan \beta }& \left(1\right)\end{array}$$\begin{array}{cc}\sigma =\frac{{\left[\frac{\left(k+1\right){\mathrm{Ma}}_f^2{\sin }^2\beta }{2+\left(k-1\right){\mathrm{Ma}}_f^2{\sin }^2\beta }\right]}^{\frac{k}{k-1}}}{{\left[\frac{2k}{k+1}{\mathrm{Ma}}_f^2{\sin }^2\beta -\frac{k-1}{k+1}\right]}^{\frac{1}{k-1}}}& \left(2\right)\end{array}$$\begin{array}{cc}{\mathrm{Ma}}_b^2=\frac{{\mathrm{Ma}}_f^2+\frac{2}{k-1}}{\frac{2k}{k-1}{\mathrm{Ma}}_f^2{\sin }^2\beta -1}+\frac{{\mathrm{Ma}}_f^2{\cos }^2\beta }{\frac{k-1}{2}{\mathrm{Ma}}_f^2{\sin }^2\beta +1}& \left(3\right)\end{array}$$\begin{array}{cc}{\sigma }_{\mathrm{inlet}}={\sigma }_F.Math.{\sigma }_1.Math.{\sigma }_2.Math.{\sigma }_n,\mathrm{wherein}n\mathrm{is}\mathrm{the}\mathrm{number}\mathrm{of}\mathrm{shock}\mathrm{waves}& \left(4\right)\end{array}$

[0050] S1.5: the flow coefficient φ.sub.i of the inlet refers to the ratio of air mass flow W.sub.ai into the inlet to air mass flow W.sub.ac through the capture area, wherein A.sub.0 represents the free flow tube area corresponding to inlet flow, A.sub.c represents the capture area which is calculated according to the geometrical relationship, and the flow coefficient is calculated according to formula 5. Given the flight altitude and Mach number, φ.sub.i=φ.sub.max is calculated through the geometrical relationship, and φ.sub.max represents the maximum flow coefficient in this state; φ.sub.i<φ.sub.max, in the subcritical state; and φ.sub.i>φ.sub.max, in the supercritical state;

[00008]$\begin{array}{cc}{\phi }_i=\frac{W_{\mathrm{ai}}}{W_{ac}}=\frac{{\rho }_0{V}_0{A}_0\_A_0}{{\rho }_0{V}_0{A}_c{A}_c}& \left(5\right)\end{array}$

[0051] S1.6: the drag of the supersonic inlet comprises external drag and external drag, wherein the internal drag (bleed drag and boundary layer suction drag) is determined by the opening degree of a bleed valve and a boundary layer suction valve, and the external drag is mainly composed of additional drag and overflow drag. The drag under the subsonic condition is mainly composed of additional drag D.sub.add, which can be calculated through the loss of momentum of the airflow before the inlet lip in the horizontal direction, and expressed by formula 6. T.sub.th, Ma.sub.th, A.sub.th and W.sub.a, th represent the throat temperature, the Mach number, the area and the flow, δ represents the total turning angle of the inlet, Ma.sub.0 represents the inlet Mach number of the inlet, A.sub.0 represents the inlet free flow tube area, and k represents the adiabatic exponent of gas;

[00009]$\begin{array}{cc}{D}_{\mathrm{add}}=\frac{W_{a,\mathrm{th}}}{{\mathrm{kMa}}_0}\left[\frac{{\mathrm{Ma}}_0}{{\mathrm{Ma}}_{\mathrm{th}}}\sqrt{\frac{T_{\mathrm{th}}}{T_0}}\left(1+{\mathrm{kMa}}_{\mathrm{th}}^2\right).Math.\cos \delta -\left(\frac{A_{\mathrm{th}}}{A_0}.Math.\cos \delta +{\mathrm{kMa}}_0^2\right)\right]& \left(6\right)\end{array}$

[0052] S1.7: under the supersonic condition, the external drag of the inlet comprises additional drag and overflow drag. when the flow coefficient of the inlet is greater than or equal to the maximum flow coefficient, the operation is under the critical or supercritical condition, and the overflow drag is 0; and when the flow coefficient of the inlet is less than the maximum flow coefficient, the operation is under the subcritical condition, the shock wave does not seal the inlet, and the overflow drag will appear. The supersonic drag D.sub.add is calculated according to formula 7, and the parameters are shown in FIG. 4. The calculation result of the above formula is small in the subcritical state, and the drag correction coefficient ΔC.sub.add can be calculated based on the Moeckel theory and expressed by formula 11, wherein P.sub.s1, P.sub.s2 and P.sub.s3 represent static pressures after shock waves, and L represents the distance of detachment of shock waves.

[00010]$\begin{array}{cc}{D}_{\mathrm{add}}=\left(P_{s1}-P_0\right)H_{e1}+\left(P_{s2}-P_0\right)H_{e2}+\left(P_{s3}-P_0\right)H_{e3}& \left(7\right)\end{array}$$\begin{array}{cc}{H}_{e2}=A_{\mathrm{th}}\sin \left({\delta }_1+{\delta }_2\right)\left(\frac{1}{\sin \left({\beta }_{2\mathrm{des}}-{\delta }_2\right)}-\frac{1}{\sin \left({\beta }_2-{\delta }_2\right)}\right)& \left(8\right)\end{array}$$\begin{array}{cc}{H}_{e3}=A_{\mathrm{th}}\sin \left({\delta }_1+{\delta }_2\right)\left(\frac{1}{\sin \left({\beta }_{2\mathrm{des}}-{\delta }_2\right)}-\frac{1}{\sin \left({\beta }_2-{\delta }_2\right)}\right)& \left(9\right)\end{array}$$\begin{array}{cc}{H}_{e1}=H_c-H_0-H_{e2}-H_{e3}& \left(10\right)\end{array}$$\begin{array}{cc}\Delta C_{\mathrm{add}}=\frac{2}{{\mathrm{kMa}}_0^2}\left[\frac{1}{2}\left(P_{s2}+P_{s3}\right)-P_{s1}\right].Math.\stackrel{\_}{L}.Math.\sin \left({\delta }_1+{\delta }_2\right)& \left(11\right)\end{array}$

[0053] The basic calculation process of the nozzle model is shown in FIG. 5, and the calculation thought is as follows:

[0054] S1.8: with the convergent-divergent nozzle as an example, calculating the critical expansion ratio π.sub.NZ,cr of the nozzle according to formula 12, where Δ.sub.μk represents the flow coefficient component of the conical nozzle, which is related to the convergent half angle α and the length of the convergent section L.sub.c of the nozzle, and β is the divergent half angle. Calculating the available expansion ratio π.sub.NZ,us according to formula 13 based on the total turbine outlet pressure and the environmental pressure, and judging the operating state (subcritical, critical and supercritical) of the nozzle; when π.sub.Nz,us≤π.sub.NZ,cr, the operation is in the subcritical or critical state; and π.sub.NZ,us>π.sub.NZ,cr, the operation is in the supercritical state.

[00011]$\begin{array}{cc}{\pi }_{\mathrm{NZ},\mathrm{cr}}=1+\frac{{\left(\frac{k+1}{2}\right)}^{\frac{k}{k-1}}-1+29{\Delta }_{\mu k}}{1+0.088\frac{\sqrt{\stackrel{\_}{A_9}-1}}{0.005+{\beta }^{1.5}}},\stackrel{\_}{A_9}=\frac{A_9}{A_8}& \left(12\right)\end{array}$$\begin{array}{cc}{\pi }_{\mathrm{NZ},\mathrm{us}}=\frac{P_7}{P_0}& \left(13\right)\end{array}$

[0055] S1.9: when the convergent-divergent nozzle is in the subcritical state, the area ratio

[00012]$\frac{A_9}{A_8}$

has no impact on the exit flow state, and the exit Mach number is less than 1; when the convergent-divergent nozzle is in the supercritical state, the area ratio

[00013]$\frac{A_9}{A_8}$

has an impact on the exit Mach number, and the exit Mach number Ma.sub.9t is obtained by iterative solution according to formula 14 (when subsonic airflow appears at the exit, Ma.sub.sub=Ma.sub.9t; and when supersonic airflow appears at the exit, Ma.sub.sup=Ma.sub.9t).

[00014]$\begin{array}{cc}\left(\frac{A_9}{A_8}\right)={\frac{1}{{\mathrm{Ma}}_{9t}}\left[\left(\frac{2}{\kappa +1}\right)\left(1+\frac{\kappa -1}{2}{\mathrm{Ma}}_{9t}^2\right)\right]}^{\frac{\left(\kappa +1\right)}{\left[2\left(\kappa -1\right)\right]}}& \left(14\right)\end{array}$

[0056] S1.10: after the area ratio of the convergent-divergent nozzle is given according to the designed expansion ratio, when the environmental back pressure changes, the nozzle will expand incompletely or excessively, forming different flow states, wherein three typical characteristic flow state points are respectively P.sub.1, P.sub.2 and P.sub.3, P.sub.8c represents the inlet total pressure of the nozzle, and the calculation formulas are as follows:

[00015]$\begin{array}{cc}{P}_1={P_{8c}\left(1+\frac{k-1}{2}{\mathrm{Ma}}_{\sup }^2\right)}^{\frac{k}{1-k}}\left({\mathrm{Ma}}_{\sup }>1\right)& \left(15\right)\end{array}$$\begin{array}{cc}{P}_3={P_{8c}\left(1+\frac{k-1}{2}{\mathrm{Ma}}_{sub}^2\right)}^{\frac{k}{1-k}}\left({\mathrm{Ma}}_{sub}<1\right)& \left(16\right)\end{array}$$\begin{array}{cc}{P}_2=P_1\left(\frac{2k}{k+1}{\mathrm{Ma}}_i^2-\frac{k-1}{k+1}\right)\left({\mathrm{Ma}}_i>1\right)& \left(17\right)\end{array}$

[0057] S1.11: after the four flow conditions of the convergent-divergent nozzle are determined, determining the flow state in the nozzle according to the back pressure condition P.sub.b, and then calculating parameters such as exit total pressure P.sub.9, static pressure P.sub.s9 and exit flow rate V.sub.9 of the nozzle.

[00016]$\begin{array}{cc}{P}_{s9}=\{\begin{array}{cc}{P}_1,& P_0<P_2\\ P_0,& P_2\le P_b\end{array}& \left(18\right)\end{array}$$\begin{array}{cc}{P}_9=\{\begin{array}{c}{P}_{8c},P_0<P_2\mathrm{or}{P}_0>P_3\\ {P_{s9}\left(1+\frac{k-1}{2}{\mathrm{Ma}}_{\sup }^2\right)}^{\frac{k}{k-1}},P_2<P_0<P_3\end{array}& \left(19\right)\end{array}$$\begin{array}{cc}{V}_9=\sqrt{2C_{p9}{T_o\left(1-\frac{P_9}{P_{s9}}\right)}^{\frac{1-k}{k}}},C_{p9}\mathrm{is}\mathrm{the}\mathrm{specific}\mathrm{heat}\mathrm{ratio}& \left(20\right)\end{array}$

[0058] S1.12: in the actual flow process of the nozzle, the actual throat flow and the actual thrust cannot reach the ideal state. The present invention calculates the flow coefficient and the thrust coefficient of the nozzle according to the known parameters by means of an engineering empirical formula, which are used for calculating the actual throat flow and the actual thrust. Formula 21 is the calculation method of the flow coefficient, wherein A.sub.7 represents the inlet area of the nozzle, and a represents the convergent half angle of the nozzle; and formula 22 is the calculation method of the thrust coefficient, wherein J.sub.c represents the impulse coefficient, J.sub.P(λ.sub.9) represents the computed impulse of the nozzle, and F.sub.N,id(π.sub.N,us) represents the ideal thrust of the nozzle.

[00017]$\begin{array}{cc}{\Phi }_N=1-0.0585\frac{\left(1+2.63\alpha \right)\alpha }{1+{\alpha }^2}\left[1-{\left(\frac{A_8}{A_7}\right)}^2\right]-0.01\left[1-e^{\left(-0.5{\alpha }^2\right)}\right]\frac{A_8}{A_7}& \left(21\right)\end{array}$$\begin{array}{cc}{C}_F=\frac{{\Phi }_N{\pi }_{N,\mathrm{us}}{J}_C{J}_P\left({\lambda }_9\right)-\frac{A_9}{A_7}}{{\Phi }_N{F}_{N,\mathrm{id}}\left({\pi }_{N,\mathrm{us}}\right)}& \left(22\right)\end{array}$

[0059] S2: establishment of component-level model of propulsion system

[0060] S2.1: FIG. 6 is a schematic diagram showing the composition of a component-level model of a typical propulsion system. Writing input/output modules of an inlet, a fan, a compressor, a combustion chamber, a high pressure turbine, a low pressure turbine, an external duct, a mixing chamber, an afterburner and a nozzle in C++ language based on gas flow and aerodynamic thermodynamic formulas.

[0061] S2.2: determining known input parameters of the model based on operating conditions and states of the model, determining the number and types of iteration variables through the common working equations, and conducting simulation calculation according to a gas process;

[0062] S2.3: the matching of the inlet, the nozzle and the engine needs to meet the flow and pressure balance, and when the engine is in a steady state or a dynamic operating state, flow, power and rotor dynamic equilibrium equations need to be satisfied simultaneously. The equilibrium equation residual of the propulsion system is represented by e. Selecting n iteration variables x based on the characteristics of the model, and conducting simultaneous solution on n common working equations:

[00018]$\begin{array}{c}\begin{array}{c}\begin{array}{c}{f}_1\left(x_0,x_1,x_2.Math.,x_n\right)=e_1\\ f_2\left(x_0,x_1,x_2.Math.,x_n\right)=e_2\end{array}\\ .Math..Math.\end{array}\\ f_n\left(x_0,x_1,x_2.Math.,x_n\right)=e_n\end{array}$

[0063] S2.4: after the input parameters of the inlet and the nozzle and the external environment variables (Mach number, flight altitude, main fuel flow, afterburner fuel flow and nozzle exit area) are determined, the problem essentially becomes a non-linear implicit equation set with unknown independent variables, which is calculated by numerical iterative algorithms, and the model is considered to obtain a reliable solution when n residual values of the common working equation approach 0.

[0064] S3: determination of variable geometric parameters of inlet and nozzle

[0065] S3.1: connecting the structure sizes (length, width and height) of the inlet as input fixed parameters to an input end, wherein the input fixed geometric parameters of a typical external compression inlet with two oblique and one direct shock waves comprise: width S, lengths L.sub.1 and L.sub.2, and height H.sub.c of the inlet, which are generally determined by the design sizes.

[0066] S3.2: connecting the rank angles δ.sub.1 and δ.sub.2, the bleed valve opening degree and the boundary layer suction opening degree of the inlet as variable parameters to the input end, wherein the parameters can be adjusted at any time in a dynamic process. The change of the rank angles will affect the geometrical relationship of the shock wave calculation of the inlet, and the mapping relationship between the bleed valve and the boundary layer suction is established according to the opening degree and the exhaust volume, which will affect the actual flow into the engine.

[0067] S3.3: connecting the inlet area A.sub.7, the length L.sub.c of the convergent section, the length L.sub.d of the divergent section, the convergent half angle α and the divergent angle β of the nozzle as input fixed parameters to the input end;

[0068] S3.4: connecting the throat area A.sub.8 and the exit area A.sub.9 of the nozzle as variable parameters to the input end;

[0069] S4: building of integrated intake/exhaust/engine computing platform of supersonic vehicle

[0070] S4.1: designing the inlet/exhaust/engine coupling component-level modeling and the iterative algorithm of supersonic vehicles by C++ programming, encapsulating the model through a dynamic link library, and introducing into a simulink module to establish a simulation platform;

[0071] S4.2: the parameters of the input end of the platform comprise structure sizes and adjustable parameters of the inlet and the nozzle, adjustable parameters of the engine model and environmental operating conditions, establishing a simulation platform of a dynamic process.

[0072] S5: analysis of calculation result of inlet/exhaust/engine coupling modeling

[0073] S5.1: with the operation condition with the maximum flight altitude and Ma=1.2 as the design operating point of the inlet, adjusting the structure parameters (S, L.sub.1, L.sub.2 and H.sub.c) of the inlet, wherein the throat area is obtained according to the maximum demand area to enable the shock waves to seal the inlet; and the structure parameters of the nozzle are determined according to the actual parameters.

[0074] S5.2: under the operation conditions of H=10 km, Ma=2, Wfa=0.9 kg/s, adjusting the secondary rank angle, and the changing rule of the thrust of the propulsion system and the installation thrust with the rank angle is shown in FIG. 7. It can be seen that with the increase of the rank angle, the thrust remains stable after a small increment, and the engine thrust remains basically unchanged when δ.sub.2=12°, indicating that the rank adjustment has a limited influence on the performance of engine components. When the rank angle increases, the installation thrust increases first, then decreases and reaches the maximum value when δ.sub.2=10°. Compared with the original state, the installation thrust is increased by 1.99%, indicating that the proper adjustment of the rank angle can significantly improve the installation thrust performance of the engine.

[0075] S5.3: under the operation conditions of H=10 km, Ma=2, Wfa=0.9 kg/s, the adjustment of the throat area A.sub.8 and the exit area A.sub.9 can significantly affect the flow state in the nozzle and the engine thrust. FIG. 8 shows the influence of the change of the throat area A.sub.8 on the thrust, and FIG. 9 shows the influence of the exit area A.sub.9 on the thrust. It can be seen that the throat area A.sub.8 has a great influence on the operating points of the engine, the thrust significantly decreases with the increase of the area, the thrust and the installation thrust reach the maximum value when A.sub.8=0.3 m.sup.2, but the installation drag of the engine decreases, indicating that the adjustment of A.sub.8 changes the state operating points of the engine and the flow demand, and reduces the overflow drag of the inlet; and under the condition of A.sub.8=0.3 m.sup.2, the adjustment of the exit area A.sub.9 can effectively increase the thrust and the installation thrust, which has a great influence, and when A.sub.9 increases to 0.55 from 0.3, the engine thrust is increased by 34%, and the installation thrust is increased by 38%. On the other hand, the adjustment of A.sub.9 has a small influence on the installation drag, indicating that the adjustment of A.sub.9 mainly affects the exit flow state of the nozzle, and has little influence on the engine operating points. Therefore, a fixed throat area A.sub.8 and a fixed area ratio

[00019]$\left(\frac{A_9}{A_8}\right)$

exist to make the performance of the propulsion system optimal.