Non-steady performance estimation method of ramjet

Abstract

The present application relates to a kind of subsonic ramjet engine unsteady state performance estimation method, including the step of calculating and determining the design point parameter of ramjet engine inlet, transition section, combustion chamber and tail nozzle;According to the flow balance relationship of design point parameter and inlet and tail nozzle, the step of calculating and determining the non-design point of ramjet engine inlet, transition section, combustion chamber and tail nozzle;Establish actual inlet operating point beta residual equation step;Actual inlet operating point beta is obtained in the whole unsteady process point iteration, obtains the actual thrust Fn of engine 实 , finally obtains the unsteady state thrust curve of engine considering volume effect.The present application can accurately and quickly calculate the thrust of subsonic ramjet engine under unsteady state condition when carrying out subsonic ramjet engine unsteady state performance design, can shorten the design iteration cycle of overall design personnel when carrying out subsonic ramjet engine overall design, and guide component design personnel to carry out detailed design.

Description

Technical Field

[0001] This invention relates to the field of aero-engines, specifically to a method for estimating the unsteady-state performance of a subsonic ramjet engine. Background Technology

[0002] Ramjet engines utilize oxygen from the air as an oxidizer, and the fuel they carry contains little or no oxidizer, resulting in a significantly increased specific impulse. Therefore, missiles powered by ramjet engines have longer ranges, lighter weights, and better maneuverability, and can achieve supersonic cruise flight throughout their flight, thus greatly enhancing their penetration capabilities. Since the 1980s, almost all countries with missile development capabilities have conducted research on ramjet propulsion technology, and various ramjet and combined ramjet engines will become the preferred power source for tactical missiles, interceptors, and cruise missiles in this century.

[0003] The performance parameters of a ramjet engine are crucial to its overall design and determine its level of sophistication. Ramjet engines undergo unsteady processes such as high-speed maneuvers during operation. Accurate design of the unsteady performance of a ramjet engine can help overall designers quickly determine whether the engine's unsteady performance parameters meet requirements, shortening the overall design iteration cycle and improving design efficiency. Summary of the Invention

[0004] The purpose of this invention is to overcome the shortcomings of the prior art and provide a method for estimating the unsteady performance of a sub-gas ramjet engine that can quickly and accurately estimate the engine's unsteady performance when designing the overall ramjet engine scheme, and facilitates the optimization and iteration of the overall engine performance.

[0005] To achieve the above objectives, the technical solution adopted by the present invention is as follows: a method for estimating the unsteady-state performance of a subsonic ramjet engine, comprising the following steps:

[0006] Step 1: Based on the given flight Mach number Ma, flight altitude H, inlet flow rate Wa, and the flow rate coefficient of the interpolation of flight Mach number Ma on the inlet interpolation characteristic diagram. Using the total pressure recovery coefficient σ, the design points for the ramjet engine's intake duct, transition section, combustion chamber, and tail nozzle are calculated and determined sequentially.

[0007] The determined intake design point parameters include: intake frontal area A. c Total temperature T at the intake and outlet of the air intake out and intake and outlet total pressure P out The determined design point parameters for the transition section include: the total pressure P at the transition section outlet. 22 Total temperature T at the outlet of the transition section 22 and the transition section inlet converted flow rate Wac 2desThe determined combustion chamber design point parameters include: combustion chamber outlet area A. 32 Total pressure P at the combustion chamber outlet 32 Combustion chamber outlet flow rate Wg 32 The determined tailpipe design point parameters include: tailpipe throat area A. 43 and tail nozzle exit area A 42 ;

[0008] Step 2: Combining the parameters of the design points for the ramjet engine's inlet, transition section, combustion chamber, and exhaust nozzle, and the flow balance relationship between the inlet and exhaust nozzle, based on publicly available standard atmospheric tables and the inlet capture flow rate Wa_c and flight Mach number Ma at non-design points... off Flight altitude H off Calculate the actual total inlet temperature T of the intake manifold. in实 and intake manifold total pressure P in实 Based on the actual inlet operating point β and the flight Mach number Ma at the non-design point. off Interpolation flow coefficient on the inlet interpolation characteristic diagram Using the total pressure recovery coefficient σ1, the non-design points of the ramjet engine's intake, transition section, combustion chamber, and exhaust nozzle are calculated and determined;

[0009] The determined inlet non-design point parameters include: inlet flow rate Wa at the non-design point. 实 Total temperature T at the intake and outlet of the air intake out实 and intake and outlet total pressure P out实 The determined non-design point parameters of the transition section include: the actual total pressure P at the outlet of the transition section. 22实 The actual total temperature T at the outlet of the transition section 22实 Actual transition section outlet flow rate Wa 22实 The determined non-design point parameters of the combustion chamber include: the actual total temperature at the combustion chamber outlet, T. 32实 Actual total pressure at the combustion chamber outlet, P 32实 Actual outlet flow rate of combustion chamber Wg 32实 The determined off-design point parameters of the exhaust nozzle include: the actual engine thrust Fn. 实 And the total pressure P8 required by the actual tail nozzle;

[0010] Step 3: Based on the total pressure P8 required for the actual tailpipe obtained in Step 2 and the known total pressure P at the actual tailpipe inlet... 41实 The residual equation is as follows:

[0011] ERR = (P 41实 -P8) / P 41实

[0012] By changing the value of the residual equation by changing the actual intake duct operating point β, the actual intake duct operating point β is iterated using Newton's method. When the absolute value of the residual is less than 0.0001, it is considered that the iteration has converged.

[0013] Step 4: Based on the residual gas coefficient α and flight Mach number Ma at the non-design point off and flight altitude H off By iterating through the actual intake manifold operating point β point by point according to the variation pattern of Time over time in step three, the actual engine thrust Fn calculated in step two can be obtained. 实 Until Time = 0 to the given time limit Time MAX The entire unsteady process was calculated, and the engine's unsteady thrust curve considering the volume effect was finally obtained.

[0014] Furthermore, step one specifically includes the following steps:

[0015] Step 11: Calculating the inlet design point: Given the flight Mach number Ma, flight altitude H, inlet flow rate Wa, and the flow rate coefficient of the flight Mach number Ma on the inlet interpolation characteristic diagram. and total pressure recovery coefficient σ;

[0016] First, calculate the total inlet temperature T of the air intake based on the standard atmospheric table and the aforementioned flight Mach number Ma and flight altitude H. in and intake manifold total pressure P in Total temperature T at the intake and exhaust outlet out Equal to the total temperature at the intake of the air duct, T in ;

[0017] And according to the formula:

[0018] P out =σP in

[0019] The total pressure P at the intake outlet was calculated. out ;

[0020] Secondly, substituting the given parameters into the flow continuity equation, we obtain:

[0021]

[0022] Wa=ρV0A

[0023]

[0024] Calculate the frontal area A of the air intake. c In the formula, K is 0.0404, which is a fixed value; q(λ) is the flow rate function, ρ is the air density, V0 is the incoming flow velocity, and A is the flow tube area;

[0025] That is, the obtained intake design point parameters include: intake frontal area A c Total temperature T at the intake and outlet of the air intake out and intake and outlet total pressure P out ;

[0026] Step 12: The steps for calculating the design point of the total pressure loss of the transition section:

[0027] Given the total pressure P at the inlet of the transition section 21 Equal to the total pressure P at the intake outlet out The total temperature T at the inlet of the transition section 21 Equal to the total temperature T at the intake duct outlet out Total temperature T at the outlet of the transition section 22 Equal to the total inlet temperature T of the transition section 21 The inlet flow rate Wa2 of the transition section is equal to the inlet flow rate Wa, and the outlet flow rate Wa of the transition section is equal to... 22 Equal to the inlet flow rate Wa of the transition section 21 And give the total pressure recovery coefficient σ of the transition section. 2des ,

[0028] Therefore, the total pressure P at the outlet of the transition section can be calculated. 22 The total pressure P at the inlet of the transition section 21 and the total pressure recovery coefficient σ of the transition section 2des Calculate the total pressure P at the outlet of the transition section. 22 The process represents the total pressure loss;

[0029] Continuing with the known parameters, calculate the equivalent flow rate Wac at the transition section inlet according to the formula. 2des ,

[0030]

[0031] That is, the obtained design point parameters for the transition section include: the total pressure P at the outlet of the transition section. 22 Total temperature T at the outlet of the transition section 22 and the transition section inlet converted flow rate Wac 2des ;

[0032] Step 13: The steps for calculating the combustion chamber design point:

[0033] It is known that the transition section outlet cross-section is the same as the combustion chamber inlet cross-section, and the total combustion chamber inlet temperature T 31 Total pressure P at the combustion chamber inlet 31 Combustion chamber cold total pressure recovery coefficient σ 3des Combustion chamber inlet flow rate Wa 31 Combustion chamber inlet Mach number 31 and combustion chamber outlet temperature T 32 ;

[0034] Substituting the known parameters into the flow continuity equation, we get:

[0035]

[0036] Wa 31 =ρ3v3A 31

[0037] Thus, the combustion chamber inlet area A can be calculated. 31 In the formula, ρ3 is the air density in the combustion chamber, and v3 is the incoming flow velocity in the combustion chamber;

[0038] Then, given the combustion efficiency η b Combined with the aforementioned combustion chamber outlet temperature T 32 This allows us to determine the fuel flow rate W in the combustion chamber. fb Then, based on the combustion chamber fuel flow rate W... fb Calculate the combustion chamber outlet flow rate Wg 32 ;

[0039] Continuing with the known parameters, calculate the converted flow rate Wac at the combustion chamber inlet according to the formula. 3des :

[0040]

[0041] Since the combustion chamber is a straight section, the combustion chamber outlet area A 32 Equal to the combustion chamber inlet area A 31 According to the total pressure P at the combustion chamber inlet 31 and the combustion chamber cold total pressure recovery coefficient σ 3des ,

[0042] Furthermore, according to the flow conservation equation, the combustion chamber inlet area A... 31 Combustion chamber inlet flow rate Wa 31 Total pressure P at the combustion chamber inlet 31 Total temperature at the combustion chamber inlet T 31 Calculate the inlet airflow velocity V of the combustion chamber. 31 ; by the combustion chamber outlet area A 32 Combustion chamber outlet flow rate Wg 32 Total pressure P at the combustion chamber outlet 32 Total temperature T at the combustion chamber outlet 32 Calculate the gas velocity V at the combustion chamber exit. 32 ; by combustion chamber inlet area A 31 Combustion chamber inlet flow rate Wa 31 Total pressure P at the combustion chamber inlet 31 Total temperature at the combustion chamber inlet T 31 Calculate the static pressure Ps at the combustion chamber inlet. 31 ;

[0043] Then, the thermal loss is calculated based on the conservation of momentum. Substituting the known parameters into the momentum conservation formula, we obtain:

[0044] Wa3V 31 +Ps 31 A 31 =Wg 32 V 32 +Ps 32 A 32

[0045] In the formula, the left side of the equation represents the impulse considering cold-state losses, and the right side represents the impulse after heating. The total pressure P at the combustion chamber outlet is obtained by iterating the momentum conservation formula as the residual. 32 And calculate the static pressure Ps at the combustion chamber outlet. 32 and the Mach number at the combustion chamber exit 32 This refers to the thermal resistance loss in the combustion chamber.

[0046] That is, the final combustion chamber design point parameters include: combustion chamber outlet area A 32 Total pressure P at the combustion chamber outlet 32 Combustion chamber outlet flow rate Wg 32 ;

[0047] Step 14: The steps for calculating the tailpipe design point:

[0048] Given that the inlet cross-sectional parameters of the exhaust nozzle are equal to those of the combustion chamber outlet cross-sectional parameters, and the total inlet temperature T of the exhaust nozzle... 41 Total pressure P at the tail nozzle inlet 41 and tail nozzle inlet flow rate Wg 41 The throat section of the exhaust nozzle is in a critical state, and the total temperature T of the exhaust nozzle throat is... 43 Equal to the total temperature T at the nozzle inlet 41 Total pressure P at the throat of the tail nozzle 43 Equal to the total pressure T at the tail nozzle inlet 41 The flow rate Wg at the throat of the tail nozzle 43 Equal to the tail nozzle inlet flow rate Wg 41 ;

[0049] Meanwhile, the inlet cross-sectional area A of the tail nozzle 41 Tail nozzle outlet cross-sectional area A 42 The cross-sectional area A of the tailpipe throat 43 The total temperature, total pressure, and flow rate of the three cross-sectional areas are the same, meaning the total temperature T at the nozzle inlet is the same. 41 Total temperature T at the tail nozzle exit 42 and the total temperature T of the tail nozzle throat 43 Equal, total pressure P at the tail nozzle inlet 41 Total pressure P at the tail nozzle exit 42and the total pressure P at the throat of the tail nozzle 43 Equal, tail nozzle inlet flow rate Wg 41 Tail nozzle outlet flow rate Wg 42 and tail nozzle throat flow rate Wg 43 equal;

[0050] Substitute the known parameters into the flow continuity equation:

[0051]

[0052] W 43 =ρv 43 A 43

[0053] Thus, the area A of the nozzle throat can be calculated. 43 Because the back pressure at the tail nozzle exit is atmospheric pressure P s0 Then, substituting the known parameters into the flow continuity equation, we get:

[0054]

[0055] Wg 42 =ρv 42 A 42

[0056] And the total static pressure relationship:

[0057] Ps0 = Ps 42 =P 42 ·π(λ)

[0058] Thus, the exhaust area A of the tail nozzle can be calculated. 43 According to the tail nozzle exit velocity V 43 Nozzle outlet flow rate Wg 43 Tail nozzle exit area A 42 and tail nozzle outlet static pressure P s42 And substituting the known inlet flow rate Wa and incoming flow velocity V0 into the formula, we get:

[0059] Fn = Wg 42 ·V 42 -Wa·V0+(P s42 -P s0 )*A 42

[0060] Thus, the engine thrust Fn can be calculated;

[0061] That is, the design point parameters of the tailpipe include: the throat area A of the tailpipe. 43 and tail nozzle exit area A 42 .

[0062] Furthermore, step two specifically includes the following steps:

[0063] Step 21: The steps for calculating the non-design point of the intake duct:

[0064] Based on publicly available standard atmospheric tables and inlet capture flow rate Wa_c at non-design points, flight Mach number Ma off Flight altitude H off Calculate the actual total inlet temperature T of the intake manifold. in实 and intake manifold total pressure P in实 Based on the actual inlet operating point β and the flight Mach number Ma at the non-design point. off Interpolation flow coefficient on the inlet interpolation characteristic diagram and total pressure recovery coefficient σ1,

[0065] Substituting the known parameters into the flow continuity equation, we get:

[0066]

[0067]

[0068] Flow coefficient and windward area A c The actual intake flow rate Wa was calculated. 实 ;

[0069] Similarly, according to the formula:

[0070] P out实 =σ1P in实

[0071] The total outlet pressure P at the non-design point was calculated. out实 Total intake outlet temperature T at non-design points out实 Equal to the total temperature at the intake of the air duct, T in实 ;

[0072] That is, the obtained inlet non-design point parameters include: inlet flow rate Wa at the non-design point. 实 Total temperature T at the intake and outlet of the air intake out实 and intake and outlet total pressure P out实 .

[0073] Step 22: The steps for calculating the non-design points of the transition section:

[0074] Based on the actual total inlet pressure P of the transition section 21实 Equal to the actual total pressure P at the intake and exhaust outlet out实 The actual total inlet temperature T of the transition section 21实 Equal to the actual total temperature T at the intake and exhaust outlet out实 Actual transition section inlet flow Wa 21Equal to the actual intake flow rate Wa 实 ;

[0075] Furthermore, based on the fact that the total pressure loss of the transition section is proportional to the square of the equivalent flow rate at the transition section inlet, and according to the equivalent flow rate Wac at the design point of the transition section inlet... 2des And the total pressure recovery coefficient σ at the design point 2des The actual total pressure recovery coefficient σ2 is calculated using the following formula:

[0076]

[0077] Based on the actual total inlet pressure P of the transition section 21实 The actual outlet total pressure P is calculated using the actual total pressure recovery coefficient σ2. 22实 The actual total temperature T at the outlet of the transition section 22实 Equal to the actual total inlet temperature T of the transition section 21实 Actual transition section outlet flow rate Wa 22实 Equal to the actual inlet flow rate Wa of the transition section 21实 ;

[0078] That is, the obtained non-design point parameters of the transition section include: the actual total pressure P at the outlet of the transition section. 22实 The actual total temperature T at the outlet of the transition section 22实 Actual transition section outlet flow rate Wa 22实 ;

[0079] Step 23: The steps for calculating the non-design point of the combustion chamber:

[0080] It is known that the transition section outlet cross-section is the same as the combustion chamber inlet cross-section, and the actual total combustion chamber inlet temperature T 31实 Actual total pressure at the combustion chamber inlet P 31实 Actual combustion chamber inlet flow rate W g31实 The actual combustion chamber outlet residual gas coefficient α and the known combustion chamber inlet area A 31 and combustion chamber outlet area A 32 ;

[0081] Substituting the known parameters into the flow continuity equation, we get:

[0082]

[0083] Wa 31实 =ρv3A 31

[0084] In the formula, Wa 31实 V represents the actual combustion chamber inlet flow rate. 3实 The actual inlet velocity of the combustion chamber is given; therefore, the actual inlet static pressure P of the combustion chamber is calculated. s31实 and imported static temperature T s31实;

[0085] The total cold pressure loss in the combustion chamber is proportional to the square of the inlet equivalent flow rate. This is based on the combustion chamber inlet equivalent flow rate Wac at the design point. 3des and the combustion chamber cold total pressure recovery coefficient σ 3des The actual total pressure recovery coefficient σ3 of the combustion chamber is calculated using the following formula:

[0086]

[0087] Then, based on the actual total pressure P at the combustion chamber inlet... 31实 The actual cold-state outlet total pressure P is calculated using the combustion chamber actual cold-state total pressure recovery coefficient σ3. 32实 The actual Mach number at the combustion chamber inlet is Ma. 31实 Actual total import pressure P 31实 The actual interpolated combustion efficiency η of the residual gas coefficient α on the combustion chamber characteristic diagram b实 According to the actual interpolated combustion efficiency η b实 The residual gas coefficient α is used to determine the actual combustion chamber outlet temperature T. 42实 ;

[0088] According to the flow conservation equation, the combustion chamber inlet area A 31 Actual combustion chamber inlet flow rate Wa 31实 Actual total pressure at the combustion chamber inlet P 31实 Actual total temperature at the combustion chamber inlet T 31实 Calculate the actual combustion chamber inlet airflow velocity V 31实 ; by the combustion chamber outlet area A 32 Actual combustion chamber outlet flow rate Wg 32实 Actual total pressure at the combustion chamber outlet, P 32实 Actual total combustion chamber outlet temperature T 32实 Calculate the actual combustion chamber outlet airflow velocity V 32实 ; by combustion chamber inlet area A 31 Actual combustion chamber inlet flow rate Wa 31实 Actual total pressure at the combustion chamber inlet P 31实 Actual total temperature at the combustion chamber inlet T 31实 Calculate the actual combustion chamber inlet static pressure Ps 31实 ;

[0089] The thermal loss is calculated based on the conservation of momentum. Substituting the known quantities into the formula yields:

[0090] Wa 31实 V 31实 +Ps 31实 A 31 =Wg 32实 V 32实 +Ps 32实A 32

[0091] In the formula, the left side of the equation represents the impulse considering cold-state losses, and the right side represents the impulse after heating. The residual iteration formula is used to obtain the actual total outlet pressure P of the combustion chamber. 32实 ;

[0092] That is, the obtained non-design point parameters of the combustion chamber include: the actual total temperature at the combustion chamber outlet, T. 32实 Actual total pressure at the combustion chamber outlet, P 32实 Actual outlet flow rate of combustion chamber Wg 32实 ;

[0093] 25) Steps for calculating the non-design point of the tailpipe:

[0094] The inlet cross-sectional parameters of the exhaust nozzle are equal to the outlet cross-sectional parameters of the combustion chamber. The actual total inlet temperature T of the exhaust nozzle is known. 41实 Actual total pressure P at the tailpipe inlet 41实 and actual tailpipe inlet flow rate Wg 41实 And the area A of the tailpipe throat at the design point. 43 Tail nozzle outlet cross-sectional area A 42 ;

[0095] The nozzle throat is first calculated under critical conditions, with the known parameter being the actual total temperature T at the nozzle throat. 43实 Actual tailpipe throat flow rate W g43实 and throat area A 43 Substituting into the flow continuity equation, we get:

[0096]

[0097] Wg 43实 =ρv8A 43实

[0098] Calculate the total pressure P8 required for the actual tailpipe.

[0099] Finally, according to the flow conservation equation, the actual tailpipe outlet flow rate W g42实 Actual total temperature T at the tailpipe outlet 42实 Tail nozzle exit area A 42 And the actual tailpipe outlet static pressure P s42 Calculate the nozzle exit velocity V9 and combine it with the actual nozzle exit flow rate W. g42实 Tail nozzle exit area A 42 And the actual tailpipe outlet static pressure P s42 And the inlet flow rate Wa and the inlet velocity V0 of the intake duct, substituting them into the formula, we get:

[0100] Fn 实 =W g42实·V 42实 -Wa·V0+(P s42实 -P s0 )*A 42

[0101] Thus, the thrust at the non-design point of the engine can be calculated, which is the actual thrust Fn of the engine. 实 ;

[0102] That is, obtaining the off-design point parameters of the exhaust nozzle includes: the actual engine thrust Fn 实 And the total pressure P8 required by the actual tail nozzle.

[0103] Furthermore, step 23 also includes measuring the total temperature T at the actual combustion chamber outlet. 32实 Actual outlet flow rate of combustion chamber (Wg) 32实 Calculation steps considering volumetric effects:

[0104] Based on the actual combustion chamber volume V 实 Current actual combustion chamber outlet flow rate W g32 Actual total enthalpy at combustion chamber outlet H 32实 Actual total pressure at the combustion chamber outlet, P 32实 And the changes in total pressure and internal energy ΔP and ΔU compared to the previous moment, calculate the total enthalpy H4 at the combustion chamber outlet considering the volume effect and the flow rate W considering the volume. g4 The formula is as follows:

[0105]

[0106] U = H 32实 -R·T 32实

[0107]

[0108] Calculate the total enthalpy H4 at the combustion chamber outlet, taking volume into account; R in the formula is the gas constant. Considering the volume effect, the actual total temperature T at the combustion chamber outlet is... 32实 The value is equal to T4, and the actual outlet flow rate of the combustion chamber is Wg. 32实 The value is equal to the flow rate W considering the volume. g4 The volume effect is calculated to reflect the impact of unsteady states on engine performance.

[0109] Furthermore, the calculation of the total pressure P8 required for the actual tailpipe in step 24 also includes the following steps:

[0110] The tail nozzle exit has two states: one is subsonic compression to the exit, with the corresponding static pressure denoted as P. s1In the first case, the material expands supersonically to the outlet and then experiences a final normal shock wave; the corresponding static pressure is denoted as P. s2 ;

[0111] Actual atmospheric pressure P s0实 Greater than P s1 At that time, the entire nozzle is in a subsonic state, and the exit static pressure is equal to atmospheric pressure. According to the static pressure P... s0 Calculate the total pressure P8 required for the tail nozzle based on the area A9, and then recalculate the throat Mach number Ma8.

[0112] Actual atmospheric pressure P s0实 In P s1 and P s2 Between the throat and the outlet, there exists a normal shock wave. The nozzle exit speed is subsonic, and the exit static pressure is equal to atmospheric pressure. Therefore, according to the static pressure P... s0 and tail nozzle exit area A 43 Calculate the actual total pressure P at the tailpipe outlet 42实 At this point, the total pressure at the nozzle throat is still P8;

[0113] Actual atmospheric pressure P s0实 Less than P s2 Then the nozzle exit is in a supersonic state, based on the total pressure P8 and the tail nozzle exit area A. 43 Calculate the actual static pressure P at the nozzle exit s9 .

[0114] Furthermore, in step four, the entire unsteady-state process is calculated without considering the combustion chamber volume effect when Time = 0, and the value of the total combustion chamber outlet temperature T4, which considers volume in step 23, is used instead of T. 32实 No need to consider the volumetric flow rate W g4 The value replaces Wg 32实 .

[0115] Furthermore, in step four, the entire unsteady-state process, when Time > 0, considers the combustion chamber volume effect, corresponding to the actual total combustion chamber outlet temperature T in step 23. 32实 The value is equal to the total outlet temperature T4 of the combustion chamber considering volume, and the actual outlet flow rate Wg of the combustion chamber. 32实 The value is equal to the flow rate W considering the volume. g4 .

[0116] The design point and non-design point refer to the specific flight conditions and engine operating states corresponding to the determined aerodynamic and thermodynamic parameters and geometric dimensions of the engine and its components during engine design. The design point is the point at which the engine's geometry is determined based on performance requirements. The engine's design point may differ from the design points of its components. Non-design points refer to the flight conditions and operating states encountered by the engine during operation that are not at the design point; these are the points where performance is determined based on the geometry.

[0117] The beneficial effects of this invention are: when designing the unsteady performance of a subsonic ramjet engine, this invention takes into account the influence of the combustion chamber volume effect, enabling accurate and rapid calculation of the thrust of the subsonic ramjet engine under unsteady conditions. In the overall design of a subsonic ramjet engine, the design of unsteady performance can shorten the design iteration cycle for overall designers and guide component designers in detailed design. Attached Figure Description

[0118] Figure 1 This is a flowchart of the design point calculation for a sub-gas turbine ramjet engine;

[0119] Figure 2 This is a flowchart of the intake duct design point calculation.

[0120] Figure 3 This is a flowchart of the combustion chamber design point calculation.

[0121] Figure 4 This is a flowchart of the non-design point calculation for a sub-gas ramjet engine;

[0122] Figure 5 This is a flowchart of the intake manifold non-design point calculation.

[0123] Figure 6 This is a flowchart of the calculation process for non-design points in the combustion chamber;

[0124] Figure 7 This is a flowchart of the calculation process for non-design points of the tail nozzle;

[0125] Figure 8 This is a flowchart of the unsteady-state performance design of a ramjet engine;

[0126] Figure 9 A calculation example of the method for estimating the unsteady performance of a ramjet engine, illustrated in the diagram. Detailed Implementation

[0127] The principles and features of the present invention are described below with reference to the accompanying drawings. The examples given are only for explaining the present invention and are not intended to limit the scope of the present invention.

[0128] To achieve the above objectives, the present invention provides the following specific embodiments:

[0129] Example 1: As Figure 1-8 As shown, a method for estimating the unsteady-state performance of a subsonic ramjet engine includes the following steps:

[0130] (1) Based on the performance requirements of the ramjet engine, calculate and determine the design points of the intake manifold, transition section, combustion chamber, and exhaust nozzle in sequence, such as Figure 1 As shown;

[0131] 11) The steps for calculating the intake duct design point are as follows: Figure 2 As shown: Given the flight Mach number Ma, flight altitude H, inlet flow rate Wa, and the flow rate coefficient of the flight Mach number Ma on the inlet interpolation characteristic diagram. And the total pressure recovery coefficient σ, the intake characteristic diagram is the flow coefficient. The relationship between the total pressure recovery coefficient σ and the flight Mach number Ma is shown in the graph.

[0132] First, calculate the total inlet temperature T of the air intake based on the standard atmospheric table and the aforementioned flight Mach number Ma and flight altitude H. in and intake manifold total pressure P in Total temperature T at the intake and exhaust outlet out Equal to the total temperature at the intake of the air duct, T in ;

[0133] And according to the formula:

[0134] P out =σP in

[0135] The total pressure P at the intake outlet was calculated. out ;

[0136] Secondly, substituting the given parameters into the flow continuity equation, we obtain:

[0137]

[0138] Wa=ρV0A

[0139]

[0140] Calculate the frontal area A of the air intake. c In the formula, K is 0.0404, which is a fixed value; q(λ) is the flow rate function, ρ is the air density, V0 is the incoming flow velocity, and A is the flow tube area;

[0141] That is, the obtained intake design point parameters include: intake frontal area A c Total temperature T at the intake and outlet of the air intake out and intake and outlet total pressure P out .

[0142] 12) Steps for calculating the design point of the total pressure loss of the transition section:

[0143] Given the total pressure P at the inlet of the transition section 21 Equal to the total pressure P at the intake outlet out The total temperature T at the inlet of the transition section 21 Equal to the total temperature T at the intake duct outlet out Total temperature T at the outlet of the transition section 22 Equal to the total inlet temperature T of the transition section 21 The inlet flow rate Wa2 of the transition section is equal to the inlet flow rate Wa, and the outlet flow rate Wa of the transition section is equal to... 22 Equal to the inlet flow rate Wa of the transition section 21 And give the total pressure recovery coefficient σ of the transition section. 2des ,

[0144] Therefore, the total pressure P at the outlet of the transition section can be calculated. 22 The total pressure P at the inlet of the transition section 21 and the total pressure recovery coefficient σ of the transition section 2des Calculate the total pressure P at the outlet of the transition section. 22 The process represents the total pressure loss;

[0145] Continuing with the known parameters, calculate the equivalent flow rate Wac at the transition section inlet according to the formula. 2des ,

[0146]

[0147] That is, the obtained design point parameters for the transition section include: the total pressure P at the outlet of the transition section. 22 Total temperature T at the outlet of the transition section 22 and the transition section inlet converted flow rate Wac 2des .

[0148] 13) The steps for calculating the combustion chamber design point are as follows: Figure 3 As shown: The known transition section outlet cross-section is the same as the combustion chamber inlet cross-section, and the total combustion chamber inlet temperature T 31 Total pressure P at the combustion chamber inlet 31 Combustion chamber cold total pressure recovery coefficient σ 3des Combustion chamber inlet flow rate Wa 31 Combustion chamber inlet Mach number 31 and combustion chamber outlet temperature T 32 ;

[0149] Substituting the known parameters into the flow continuity equation, we get:

[0150]

[0151] Wa 31 =ρ3v3A 31

[0152] Thus, the combustion chamber inlet area A can be calculated. 31 In the formula, ρ3 is the air density in the combustion chamber, and v3 is the incoming flow velocity in the combustion chamber;

[0153] Then, given the combustion efficiency η b Combined with the aforementioned combustion chamber outlet temperature T 32 This allows us to determine the fuel flow rate W in the combustion chamber. fb Then, based on the combustion chamber fuel flow rate W... fb Calculate the combustion chamber outlet flow rate Wg 32 ;

[0154] Continuing with the known parameters, calculate the converted flow rate Wac at the combustion chamber inlet according to the formula. 3des :

[0155]

[0156] Since the combustion chamber is a straight section, the combustion chamber outlet area A 32 Equal to the combustion chamber inlet area A 31 According to the total pressure P at the combustion chamber inlet 31 and the combustion chamber cold total pressure recovery coefficient σ 3des ,

[0157] Furthermore, according to the flow conservation equation, the combustion chamber inlet area A... 31 Combustion chamber inlet flow rate Wa 31 Total pressure P at the combustion chamber inlet 31 Total temperature at the combustion chamber inlet T 31 Calculate the inlet airflow velocity V of the combustion chamber. 31 ; by the combustion chamber outlet area A 32 Combustion chamber outlet flow rate Wg 32 Total pressure P at the combustion chamber outlet 32 Total temperature T at the combustion chamber outlet 32 Calculate the gas velocity V at the combustion chamber exit. 32 ; by combustion chamber inlet area A 31 Combustion chamber inlet flow rate Wa 31 Total pressure P at the combustion chamber inlet 31 Total temperature at the combustion chamber inlet T 31 Calculate the static pressure Ps at the combustion chamber inlet. 31 ;

[0158] Then, the thermal loss is calculated based on the conservation of momentum. Substituting the known parameters into the momentum conservation formula, we obtain:

[0159] Wa3V 31 +Ps 31 A 31 =Wg32 V 32 +Ps 32 A 32

[0160] In the formula, the left side of the equation represents the impulse considering cold-state losses, and the right side represents the impulse after heating. The total pressure P at the combustion chamber outlet is obtained by iterating the momentum conservation formula as the residual. 32 And calculate the static pressure Ps at the combustion chamber outlet. 32 and the Mach number at the combustion chamber exit 32 This refers to the thermal resistance loss in the combustion chamber.

[0161] That is, the final combustion chamber design point parameters include: combustion chamber outlet area A 32 Total pressure P at the combustion chamber outlet 32 Combustion chamber outlet flow rate Wg 32 .

[0162] 14) Steps for calculating the design point of the tailpipe: Given that the inlet cross-sectional parameters of the tailpipe are equal to the outlet cross-sectional parameters of the combustion chamber, and the total inlet temperature T of the tailpipe. 41 Total pressure P at the tail nozzle inlet 41 and tail nozzle inlet flow rate Wg 41 The throat section of the exhaust nozzle is in a critical state, and the total temperature T of the exhaust nozzle throat is... 43 Equal to the total temperature T at the nozzle inlet 41 Total pressure P at the throat of the tail nozzle 43 Equal to the total pressure T at the tail nozzle inlet 41 The flow rate Wg at the throat of the tail nozzle 43 Equal to the tail nozzle inlet flow rate Wg 41 ;

[0163] Meanwhile, the inlet cross-sectional area A of the tail nozzle 41 Tail nozzle outlet cross-sectional area A 42 The cross-sectional area A of the tailpipe throat 43 The total temperature, total pressure, and flow rate of the three cross-sectional areas are the same, meaning the total temperature T at the nozzle inlet is the same. 41 Total temperature T at the tail nozzle exit 42 and the total temperature T of the tail nozzle throat 43 Equal, total pressure P at the tail nozzle inlet 41 Total pressure P at the tail nozzle exit 42 and the total pressure P at the throat of the tail nozzle 43 Equal, tail nozzle inlet flow rate Wg 41 Tail nozzle outlet flow rate Wg 42 and tail nozzle throat flow rate Wg 43 equal;

[0164] Substitute the known parameters into the flow continuity equation:

[0165]

[0166] W 43 =ρv 43 A 43

[0167] Thus, the area A of the nozzle throat can be calculated. 43 Because the back pressure at the tail nozzle exit is atmospheric pressure P s0 ,

[0168] Substituting the known parameters into the flow continuity equation, we obtain:

[0169]

[0170] Wg 42 =ρv 42 A 42

[0171] And the total static pressure relationship:

[0172] Ps0 = Ps 42 =P 42 ·π(λ)

[0173] Thus, the exhaust area A of the tail nozzle can be calculated. 43 According to the tail nozzle exit velocity V 43 Nozzle outlet flow rate Wg 43 Tail nozzle exit area A 42 and tail nozzle outlet static pressure P s42 And substituting the known inlet flow rate Wa and incoming flow velocity V0 into the formula, we get:

[0174] Fn = Wg 42 ·V 42 -Wa·V0+(P s42 -P s0 )*A 42

[0175] Thus, the engine thrust Fn can be calculated;

[0176] That is, the design point parameters of the tailpipe include: the throat area A of the tailpipe. 43 and tail nozzle exit area A 42 .

[0177] (2) Based on the parameters of the design points of the ramjet engine's intake duct, transition section, combustion chamber, and exhaust nozzle obtained in step (1), and according to the flow balance relationship between the intake duct and the exhaust nozzle, calculate and determine the non-design points of the ramjet engine's performance, such as... Figure 4 As shown;

[0178] 21) The steps for calculating the non-design point of the intake duct are as follows: Figure 5 As shown: Based on publicly available standard atmospheric tables and inlet capture flow rate Wa_c and flight Mach number Ma at non-design points... off Flight altitude H off Calculate the actual total inlet temperature T of the intake manifold. in实 and intake manifold total pressure P in实 Based on the actual inlet operating point β and the flight Mach number Ma at the non-design point. off Interpolation flow coefficient on the inlet interpolation characteristic diagram and total pressure recovery coefficient σ1,

[0179] Substituting the known parameters into the flow continuity equation, we get:

[0180]

[0181]

[0182] Flow coefficient and windward area A c The actual intake flow rate Wa was calculated. 实 ;

[0183] Similarly, according to the formula:

[0184] P out实 =σ1P in实

[0185] The total outlet pressure P at the non-design point was calculated. out实 Total intake outlet temperature T at non-design points out实 Equal to the total temperature at the intake of the air duct, T in实 .

[0186] That is, the obtained inlet non-design point parameters include: inlet flow rate Wa at the non-design point. 实 Total temperature T at the intake and outlet of the air intake out实 and intake and outlet total pressure P out实 .

[0187] 22) Steps for calculating the non-design point of the transition section: Based on the actual total inlet pressure P of the transition section 21实 Equal to the actual total pressure P at the intake and exhaust outlet out实 The actual total inlet temperature T of the transition section 21实 Equal to the actual total temperature T at the intake and exhaust outlet out实 Actual transition section inlet flow Wa 21 Equal to the actual intake flow rate Wa 实 ;

[0188] Furthermore, based on the fact that the total pressure loss of the transition section is proportional to the square of the equivalent flow rate at the transition section inlet, and according to the equivalent flow rate Wac at the design point of the transition section inlet...2des And the total pressure recovery coefficient σ at the design point 2des The actual total pressure recovery coefficient σ2 is calculated using the following formula.

[0189]

[0190] Based on the actual total inlet pressure P of the transition section 21实 The actual outlet total pressure P is calculated using the actual total pressure recovery coefficient σ2. 22实 The actual total temperature T at the outlet of the transition section 22实 Equal to the actual total inlet temperature T of the transition section 21实 Actual transition section outlet flow rate Wa 22实 Equal to the actual inlet flow rate Wa of the transition section 21实 ;

[0191] That is, the obtained non-design point parameters of the transition section include: the actual total pressure P at the outlet of the transition section. 22实 The actual total temperature T at the outlet of the transition section 22实 Actual transition section outlet flow rate Wa 22实 .

[0192] 23) The steps for calculating the non-design point of the combustion chamber are as follows: Figure 6 As shown: The known transition section outlet cross-section is the same as the combustion chamber inlet cross-section, and the actual total combustion chamber inlet temperature T 31实 Actual total pressure at the combustion chamber inlet P 31实 Actual combustion chamber inlet flow rate W g31实 The actual combustion chamber outlet residual gas coefficient α and the known combustion chamber inlet area A 31 and combustion chamber outlet area A 32 ;

[0193] Substituting the known parameters into the flow continuity equation, we get:

[0194]

[0195] Wa 31实 =ρv3A 31

[0196] In the formula, Wa 31实 V represents the actual combustion chamber inlet flow rate. 3实 The actual inlet velocity of the combustion chamber is given; therefore, the actual inlet static pressure P of the combustion chamber is calculated. s31实 and imported static temperature T s31实 ;

[0197] The total cold pressure loss in the combustion chamber is proportional to the square of the inlet equivalent flow rate. This is based on the combustion chamber inlet equivalent flow rate Wac at the design point. 3des and the combustion chamber cold total pressure recovery coefficient σ 3desThe actual total pressure recovery coefficient σ3 of the combustion chamber is calculated using the following formula:

[0198]

[0199] Then, based on the actual total pressure P at the combustion chamber inlet... 31实 The actual cold-state outlet total pressure P is calculated using the combustion chamber actual cold-state total pressure recovery coefficient σ3. 32实 The actual Mach number at the combustion chamber inlet is Ma. 31实 Actual total import pressure P 31实 The actual interpolated combustion efficiency η of the residual gas coefficient α on the combustion chamber characteristic diagram b实 According to the actual interpolated combustion efficiency η b实 The residual gas coefficient α is used to determine the actual combustion chamber outlet temperature T. 42实 ;

[0200] According to the flow conservation equation, the combustion chamber inlet area A 31 Actual combustion chamber inlet flow rate Wa 31实 Actual total pressure at the combustion chamber inlet P 31实 Actual total temperature at the combustion chamber inlet T 31实 Calculate the actual combustion chamber inlet airflow velocity V 31实 ; by the combustion chamber outlet area A 32 Actual combustion chamber outlet flow rate Wg 32实 Actual total pressure at the combustion chamber outlet, P 32实 Actual total combustion chamber outlet temperature T 32实 Calculate the actual combustion chamber outlet airflow velocity V 32实 ; by combustion chamber inlet area A 31 Actual combustion chamber inlet flow rate Wa 31实 Actual total pressure at the combustion chamber inlet P 31实 Actual total temperature at the combustion chamber inlet T 31实 Calculate the actual combustion chamber inlet static pressure Ps 31实 ;

[0201] The thermal loss is calculated based on the conservation of momentum. Substituting the known quantities into the formula yields:

[0202] Wa 31实 V 31实 +Ps 31实 A 31 =Wg 32实 V 32实 +Ps 32实 A 32

[0203] In the formula, the left side of the equation represents the impulse considering cold-state losses, and the right side represents the impulse after heating. The residual iteration formula is used to obtain the actual total outlet pressure P of the combustion chamber. 32实 ;

[0204] Based on the actual combustion chamber volume V 实 Current actual combustion chamber outlet flow rate W g32 Actual total enthalpy at combustion chamber outlet H 32实 Actual total pressure at the combustion chamber outlet, P 32实 And the changes in total pressure and internal energy ΔP and ΔU compared to the previous moment, calculate the total enthalpy H4 at the combustion chamber outlet considering the volume effect and the flow rate W considering the volume. g4 The formula is as follows:

[0205]

[0206] U = H 32实 -R·T 32实

[0207]

[0208] Calculate the total enthalpy H4 at the combustion chamber outlet after considering the volume;

[0209] In the formula, R is the gas constant. Considering the volume effect, the actual total temperature T at the combustion chamber outlet is... 32实 The value is equal to T4, and the actual outlet flow rate of the combustion chamber is Wg. 32实 The value is equal to the flow rate W considering the volume. g4 The volume effect is calculated to reflect the impact of unsteady states on engine performance;

[0210] That is, the obtained non-design point parameters of the combustion chamber include: the actual total temperature at the combustion chamber outlet, T. 32实 Actual total pressure at the combustion chamber outlet, P 32实 Actual outlet flow rate of combustion chamber Wg 32实 .

[0211] 24) The steps for calculating the non-design point of the tailpipe are as follows: Figure 7 As shown:

[0212] The inlet cross-sectional parameters of the exhaust nozzle are equal to the outlet cross-sectional parameters of the combustion chamber. The actual total inlet temperature T of the exhaust nozzle is known. 41实 Actual total pressure P at the tailpipe inlet 41实 and actual tailpipe inlet flow rate Wg 41实 And the area A of the tailpipe throat at the design point. 43 Tail nozzle outlet cross-sectional area A 42 ;

[0213] The nozzle throat is first calculated under critical conditions, with the known parameter being the actual total temperature T at the nozzle throat. 43实 Actual tailpipe throat flow rate W g43实 and throat area A43 Substituting into the flow continuity equation, we get:

[0214]

[0215] Wg 43实 =ρv8A 43实

[0216] Calculate the total pressure P8 required for the tail nozzle.

[0217] The tail nozzle outlet has two key states: one is subsonic compression to the outlet, with the corresponding static pressure denoted as P. s1 In the first case, the material expands supersonically to the outlet and then experiences a final normal shock wave; the corresponding static pressure is denoted as P. s2 ;

[0218] Actual atmospheric pressure P s0实 Greater than P s1 At that time, the entire nozzle is in a subsonic state, and the exit static pressure is equal to atmospheric pressure. According to the static pressure P... s0 Calculate the total pressure P8 required for the tail nozzle based on the area A9, and then recalculate the throat Mach number Ma8.

[0219] Actual atmospheric pressure P s0实 In P s1 and P s2 Between the throat and the outlet, there exists a normal shock wave. The nozzle exit speed is subsonic, and the exit static pressure is equal to atmospheric pressure. Therefore, according to the static pressure P... s0 and tail nozzle exit area A 43 Calculate the actual total pressure P at the tailpipe outlet 42实 At this point, the total pressure at the nozzle throat is still P8;

[0220] Actual atmospheric pressure P s0实 Less than P s2 Then the nozzle exit is in a supersonic state, based on the total pressure P8 and the tail nozzle exit area A. 43 Calculate the actual static pressure P at the nozzle exit s9 ;

[0221] Finally, according to the flow conservation equation, the actual tailpipe outlet flow rate W g42实 Actual total temperature T at the tailpipe outlet 42实 Tail nozzle exit area A 42 And the actual tailpipe outlet static pressure P s42 Calculate the nozzle exit velocity V9 and combine it with the actual nozzle exit flow rate W. g42实 Tail nozzle exit area A 42 And the actual tailpipe outlet static pressure P s42 And the inlet flow rate Wa and the inlet velocity V0 of the intake duct, substituting them into the formula, we get:

[0222] Fn 实 =W g42实 ·V 42实 -Wa·V0+(P s42实 -P s0 )*A 42

[0223] Thus, the thrust at the non-design point of the engine can be calculated, which is the actual thrust Fn of the engine. 实 ;

[0224] That is, obtaining the off-design point parameters of the exhaust nozzle includes: the actual engine thrust Fn 实 And the total pressure P8 required by the actual tail nozzle.

[0225] (3) The total pressure P8 required for the tail nozzle obtained from step 24) and the actual total pressure P at the tail nozzle inlet 41实 The residual equation is as follows:

[0226] ERR = (P 41实 -P8) / P 41实

[0227] The value of the residual equation can be changed by changing the actual intake duct operating point β described in step 21). The actual intake duct operating point β is iterated using Newton's method. When the absolute value of the residual is less than 0.0001, it is considered that the iteration has converged.

[0228] (4) Figure 8 As shown, based on the residual gas coefficient α at the non-design point and the flight Mach number Ma... off and flight altitude H off By iterating through the actual intake working point β in step (3) according to the variation law of Time, the actual engine thrust Fn calculated in step (2) is obtained. 实 Until the entire unsteady-state process is calculated, at Time = 0, the combustion chamber volume effect is not considered, and the value of the total combustion chamber outlet temperature T4, which considers volume in step 23, is used to replace T. 32实 No need to consider the volumetric flow rate W g4 The value replaces Wg 32实 When Time > 0, the combustion chamber volume effect is considered, corresponding to the actual total combustion chamber outlet temperature T mentioned in step 23. 32实 The value is equal to the total outlet temperature T4 of the combustion chamber considering volume, and the actual outlet flow rate Wg of the combustion chamber. 32实 The value is equal to the flow rate W considering the volume. g4 The calculation continues until Time reaches the given time limit. MAX Finally, the unsteady thrust curve of the engine considering the volume effect is obtained.

[0229] The specific calculation example is as follows:

[0230] A method for designing the unsteady-state performance of a subsonic ramjet engine includes the following calculation steps:

[0231] (1) This specific implementation takes the design of the unsteady-state performance of a certain subsonic ramjet engine as an example. When calculating the inlet design point, flight conditions and inlet characteristic parameters are required, including flight Mach number Ma = 4.0, flight altitude H = 21500m, inlet flow rate Wa = 10.0kg / s, total pressure recovery coefficient σ = 0.524, and flow coefficient. Based on the above parameters, atmospheric conditions, and the flow continuity equation... The sum of Wa = ρvA can be used to calculate the air intake frontal area A = 0.12225m². 2 The total outlet pressure P2 = 344339.31 Pa and the total outlet temperature T2 = 916.23 K.

[0232] (2) Based on the above parameters, the total pressure recovery coefficient σ at the design point of the transition section is given. 2des =0.9, so the total pressure at the transition section outlet is P3 = 300608.22 Pa and the total temperature is T3 = 916.23 K.

[0233] (3) By giving the combustion chamber cold total pressure recovery coefficient σ 3des =0.93, combustion efficiency η b =0.92, inlet Mach number Ma3 = 0.1, outlet total temperature T4 = 2000K, based on the conservation of momentum and energy, the outlet total pressure P4 = 276576.41Pa, and the combustion chamber area A3 = 0.14805m². 2 Oil supply W fb =0.37660 kg / s, and the exit Mach number is Ma4 = 0.1733.

[0234] (4) Determine the tail nozzle exit back pressure P based on flight altitude s0 =4325.2 Pa, and the throat area A8 = 0.04292 m² is calculated from the critical throat condition. 2 The nozzle exit area A9 is calculated to be 0.30138 m² based on the full expansion condition. 2 The exit Mach number is Ma9 = 3.2495, and the exit velocity is V9 = 1766.9 m / s. Based on the flight Mach number and altitude, the incoming flow velocity is calculated to be V0 = 1184.3 m / s. Using the thrust calculation formula, Fn = 614.94 kgf.

[0235] (5) Based on the calculation results at the design point, the fuel supply W for the unsteady-state process of the ramjet engine is given. fbThe variation of flight Mach number Ma and flight altitude H with time (Time) was investigated. Flight altitude H = 21500m, flight Mach number Ma = 4.0, which remained constant. Fuel supply increased linearly from 0.2 kg / s to 0.3766 kg / s within 3.6 seconds. The calculated time step was 0.025 s. Combustion chamber volume V = 1 m³. 3 The non-design point model established using this method is used to design the engine's unsteady-state performance. The thrust variation law is as follows: Figure 9 As shown, the thrust of the ramjet engine changes from 316.22 kgf to 614.94 kgf.

[0236] The above description is only a preferred embodiment of the present invention and is not intended to limit the present invention. Any modifications, equivalent substitutions, improvements, etc., made within the spirit and principles of the present invention should be included within the protection scope of the present invention.

Claims

1. A method of estimating unsteady performance of a scramjet engine, characterized by, It comprises the following steps: Step one, according to the given flight Mach number Ma, flight height H, inlet inlet flow Wa, and flight Mach number Ma on the interpolation characteristic diagram of the interpolation flow coefficient φ and total pressure recovery coefficient σ, in turn calculate and determine the ramjet inlet, transition section, combustion chamber and tail nozzle design point; The determined inlet design point parameters include: inlet windward area A c , inlet exit total temperature T out , and inlet exit total pressure P out ; the determined transition section design point parameters include: transition section exit total pressure P 22 , transition section exit total temperature T 22 , and transition section inlet equivalent flow Wac 2des ; the determined combustion chamber design point parameters include: combustion chamber exit area A 32 , combustion chamber exit total pressure P 32 , and combustion chamber exit flow Wg 32 ; and the determined tail nozzle design point parameters include: tail nozzle throat area A 43 , and tail nozzle exit area A 42 . Step 2: Combining the parameters of the design points for the ramjet engine's inlet, transition section, combustion chamber, and exhaust nozzle, and the flow balance relationship between the inlet and exhaust nozzle, based on publicly available standard atmospheric tables and the inlet capture flow rate Wa_c and flight Mach number Ma at non-design points... off Flight altitude H off Calculate the actual total inlet temperature T of the intake manifold. in实 and intake manifold total pressure P in实 Based on the actual inlet operating point β and the flight Mach number Ma at the non-design point. off The non-design points of the ramjet engine's intake duct, transition section, combustion chamber, and tail nozzle are calculated and determined using the interpolation flow coefficient φ1 and total pressure recovery coefficient σ1 on the intake interpolation characteristic diagram. The determined non-design point parameters of the inlet passage include: non-design point inlet passage flow Wa 实 , inlet passage outlet total temperature T out实 , and inlet passage outlet total pressure P out实 ; the determined non-design point parameters of the transition section include: actual transition section outlet total pressure P 22实 , actual transition section outlet total temperature T 22实 , and actual transition section outlet flow Wa 22实 ; the determined non-design point parameters of the combustion chamber include: actual combustion chamber outlet total temperature T 32实 , actual combustion chamber outlet total pressure P 32实 , and combustion chamber actual outlet flow Wg 32实 ; and the determined non-design point parameters of the tail nozzle include: engine actual thrust Fn 实 and actual tail nozzle required total pressure P8. Also included is the actual combustion chamber outlet total temperature T 32实 , the combustion chamber actual outlet flow rate Wg 32实 Calculation step considering the volume effect: According to the actual combustion chamber volume V 实 , the current actual combustion chamber outlet flow W g32 , the actual combustion chamber outlet total enthalpy H 32实 , the actual combustion chamber outlet total pressure P 32实 , and the changes of total pressure and internal energy compared with the previous moment ΔP and ΔU, the combustion chamber outlet total enthalpy H4 after considering the volume effect and the flow W considering the volume are calculated g4 , and the formula is as follows: ， According to the total enthalpy H4 at the outlet of the combustion chamber calculated by considering the volume, the total temperature T4 at the outlet of the combustion chamber is calculated by considering the volume; R in the formula is the gas constant, and the actual total temperature T 32实 at the outlet of the combustion chamber is equal to T4, the actual flow rate Wg 32实 at the outlet of the combustion chamber is equal to the flow rate W g4 considering the volume, and the calculation of the volume effect is used to reflect the influence of the non-steady state on the engine performance. Step three, the total pressure P8 required for the actual nozzle is obtained according to the actual nozzle inlet total pressure P 41实 The residual equation is constructed as shown in the following formula: ， By changing the actual inlet operating point β, the value of the residual equation is changed, and the Newton method is used to iterate the actual inlet operating point β, and when the absolute value of the residual is less than 0.0001, it is considered that the iteration converges; Step four, according to the non-design point residual coefficient α, flight Mach number Ma off and flight height H off The change rule of time Time, the actual inlet working point β is iterated point by point according to step three, so as to obtain the actual engine thrust Fn calculated in step two 实 , until Time=0 to the given time limit Time MAX The whole non-steady state process is calculated, and the engine non-steady state thrust curve considering volume effect is finally obtained.

2. The method of claim 1, wherein, Said step one specifically comprises the following steps: Step 11: the step of calculating the inlet design point: given flight Mach number Ma, flight height H, inlet inlet flow Wa, and flight Mach number Ma on the interpolation characteristic diagram of the interpolation flow coefficient φ and total pressure recovery coefficient σ; First, the total temperature T at the inlet of the inlet duct is calculated from the standard atmosphere table and the flight Mach number Ma, flight height H in and the total pressure P at the inlet of the inlet duct in The total temperature T at the outlet of the inlet duct out is equal to the total temperature T at the inlet of the inlet duct in ; And according to the formula: P out =σP in The total pressure P at the exit of the inlet duct was calculated out ; Secondly, the given parameters are brought into the flow continuity equation, and the following is obtained: ， The windward area A of the intake passage is calculated c where K is 0.0404, a fixed value; q(λ) is a flow function; p is the air density; V0 is the incoming flow velocity; and A is the flow tube area. That is, the obtained intake passage design point parameters include: an intake passage windward area A c , an intake passage outlet total temperature T out , and an intake passage outlet total pressure P out ; Step 12: the step of calculating the design point of the transition section total pressure loss: The total pressure at the inlet of the transition section P 21 is equal to the total pressure at the outlet of the inlet duct P out The total temperature at the inlet of the transition section T 21 is equal to the total temperature at the outlet of the inlet duct T out The total temperature at the outlet of the transition section T 22 is equal to the total temperature at the inlet of the transition section T 21 The flow rate at the inlet of the transition section Wa2 is equal to the flow rate at the inlet duct Wa, the flow rate at the outlet of the transition section Wa 22 is equal to the flow rate at the inlet of the transition section Wa 21 The total pressure recovery coefficient of the transition section σ 2des is given, Therefore, the total pressure P at the outlet of the transition section can be calculated. 22 The total pressure P at the inlet of the transition section 21 and the total pressure recovery coefficient σ of the transition section 2des Calculate the total pressure P at the outlet of the transition section. 22 The process represents the total pressure loss; Continuing with the known parameters, the transition inlet converted flow rate Wac is calculated according to the formula 2des , ， That is, the obtained transition section design point parameters include: transition section outlet total pressure P 22 , transition section outlet total temperature T 22 , and transition section inlet equivalent flow rate Wac 2des ; Step 13: the step of calculating the combustion chamber design point: It is known that the transition section outlet section is the combustion chamber inlet section, and the combustion chamber inlet total temperature T 31 , the combustion chamber inlet total pressure P 31 , the combustion chamber cold-state total pressure recovery coefficient σ 3des , the combustion chamber inlet flow rate Wa 31 , the combustion chamber inlet Mach number Ma 31 , and the combustion chamber outlet temperature T 32 ; The known parameters are substituted into the flow continuity equation to obtain: ， Thus, the inlet area A of the combustion chamber is calculated 31 wherein p3 is the air density in the combustion chamber and v3 is the flow velocity in the combustion chamber. Then, given the combustion efficiency η b , in combination with the combustion chamber outlet temperature T 32 , the combustion chamber fuel flow rate W fb is determined, and then, according to the combustion chamber fuel flow rate W fb , the combustion chamber outlet flow rate Wg 32 is calculated. Continuing with the known parameters, the combustion chamber inlet flow rate Wac is calculated according to the formula 3des : ， Since the combustion chamber is a constant area duct, the combustion chamber exit area A 32 is equal to the combustion chamber inlet area A 31 ; and the combustion chamber inlet total pressure P 31 and the combustion chamber cold flow total pressure recovery coefficient σ 3des , Furthermore, according to the flow conservation equation, the combustion chamber inlet area A... 31 Combustion chamber inlet flow rate Wa 31 Total pressure P at the combustion chamber inlet 31 Total temperature at the combustion chamber inlet T 31 Calculate the inlet airflow velocity V of the combustion chamber 31 ; by the combustion chamber outlet area A 32 Combustion chamber outlet flow rate Wg 32 Total pressure P at the combustion chamber outlet 32 Total temperature T at the combustion chamber outlet 32 Calculate the gas flow velocity V at the combustion chamber exit. 32 ; by the combustion chamber inlet area A 31 Combustion chamber inlet flow rate Wa 31 Total pressure P at the combustion chamber inlet 31 Total temperature at the combustion chamber inlet T 31 Calculate the static pressure Ps at the combustion chamber inlet. 31 ; Then, according to the momentum conservation calculation of thermal state loss, the known parameters are substituted into the momentum conservation formula to obtain: ， In the formula, the left side of the equal sign is the impulse considering the cold loss, and the right side is the impulse after heating. The total pressure P at the outlet of the combustion chamber is obtained by iteration of the momentum conservation formula as the residual 32 The static pressure Ps at the outlet of the combustion chamber is calculated 32 The Mach number Ma at the outlet of the combustion chamber is calculated 32 That is, the thermal resistance loss of the combustion chamber; That is, the final obtained combustion chamber design point parameters include: combustion chamber outlet area A 32 , combustion chamber outlet total pressure P 32 , and combustion chamber outlet flow rate Wg 32 ; Step 14: the step of calculating the tail nozzle design point: Given that the inlet cross-sectional parameters of the exhaust nozzle are equal to those of the combustion chamber outlet cross-sectional parameters, and the total inlet temperature T of the exhaust nozzle... 41 Total pressure P at the tail nozzle inlet 41 and tail nozzle inlet flow rate Wg 41 The throat section of the exhaust nozzle is in a critical state, and the total temperature T of the exhaust nozzle throat is... 43 Equal to the total temperature T at the nozzle inlet 41 Total pressure P at the throat of the tail nozzle 43 Equal to the total pressure T at the tail nozzle inlet 41 The flow rate Wg at the throat of the tail nozzle 43 Equal to the tail nozzle inlet flow rate Wg 41 ; At the same time, the tail nozzle inlet cross-sectional area A 41 , the tail nozzle outlet cross-sectional area A 42 , and the tail nozzle throat cross-sectional area A 43 The total temperature, total pressure and flow rate of the three cross-sectional areas are the same, that is, the tail nozzle inlet total temperature T 41 , the tail nozzle outlet total temperature T 42 , and the tail nozzle throat total temperature T 43 are equal, the tail nozzle inlet total pressure P 41 , the tail nozzle outlet total pressure P 42 , and the tail nozzle throat total pressure P 43 are equal, the tail nozzle inlet flow rate Wg 41 , the tail nozzle outlet flow rate Wg 42 , and the tail nozzle throat flow rate Wg 43 are equal; The known parameters are brought into the flow continuity equation: ， The throat area A of the tail nozzle is calculated 43 Since the back pressure at the tail nozzle outlet is atmospheric pressure P s0 The known parameters are then continuously brought into the flow equation, resulting in: ， And the total static pressure relationship: ， The nozzle exit area A is calculated 43 from the nozzle exit velocity V 43 , the nozzle exit flow rate Wg 43 , the nozzle exit area A 42 , and the nozzle exit static pressure P s42 and the known inlet duct inlet flow rate Wa, the incoming flow velocity V0, by substitution into the formula ， So as to calculate the engine thrust Fn; That is, obtaining the nozzle design point parameters includes: nozzle throat area A 43 and nozzle exit area A 42 .

3. The method of claim 1, wherein Said step two specifically comprises the following steps: Step 21: the step of calculating the non-design point of the inlet: Based on publicly available standard atmospheric tables and inlet capture flow rate Wa_c at non-design points, flight Mach number Ma off Flight altitude H off Calculate the actual total inlet temperature T of the intake manifold. in实 and intake manifold total pressure P in实 Based on the actual inlet operating point β and the flight Mach number Ma at the non-design point. off The interpolated flow rate coefficient φ1 and total pressure recovery coefficient σ1 on the inlet interpolation characteristic diagram The known parameters are substituted into the flow continuity equation to obtain: ， Flow coefficient φ1 and windward area A c The flow Wa of the actual intake is calculated 实 ; Similarly, according to the formula: P out实 =σ1P in实 The total pressure at the exit of the inlet duct P out实 at the non-design point is calculated to be equal to the total temperature at the inlet of the inlet duct T out实 at the non-design point is calculated to be equal to the total temperature at the inlet of the inlet duct T in实 ; That is, the obtained intake passage off-design point parameters include: intake passage flow rate Wa 实 , intake passage exit total temperature T out实 , and intake passage exit total pressure P out实 ; Step 22: the step of calculating the non-design point of the transition section: According to the actual transition section inlet total pressure P 21实 is equal to the actual inlet duct outlet total pressure P out实 , the actual transition section inlet total temperature T 21实 is equal to the actual inlet duct outlet total temperature T out实 , the actual transition section inlet flow rate Wa 21 is equal to the actual inlet duct flow rate Wa 实 ; According to the fact that the total pressure loss of the transition section is proportional to the square of the equivalent flow rate at the inlet of the transition section, and according to the equivalent flow rate Wac at the inlet of the transition section of the design point 2des , and the total pressure recovery coefficient σ of the design point 2des , the actual total pressure recovery coefficient σ2 is calculated according to the following formula: ， Based on the actual total inlet pressure P of the transition section 21实 The actual outlet total pressure P is calculated using the actual total pressure recovery coefficient σ2. 22实 The actual total temperature T at the outlet of the transition section 22实 Equal to the actual total inlet temperature T of the transition section 21实 Actual transition section outlet flow rate Wa 22实 Equal to the actual transition section inlet flow Wa 21实 ; That is, the obtained transition section off-design parameters include: actual transition section outlet total pressure P 22实 , actual transition section outlet total temperature T 22实 , actual transition section outlet flow rate Wa 22实 ; Step 23: the step of calculating the non-design point of the combustion chamber: The transition section outlet cross section is known to be the combustion chamber inlet cross section, and the actual combustion chamber inlet total temperature T 31实 , the actual combustion chamber inlet total pressure P 31实 , the actual combustion chamber inlet flow rate W g31实 , the actual combustion chamber outlet residual coefficient a and the known combustion chamber inlet area A 31 and the combustion chamber outlet area A 32 ; The known parameters are substituted into the flow continuity equation to obtain: ， wherein Wa 31实 is the actual combustor inlet flow rate, V 3实 is the actual combustor inlet flow velocity; whereby the actual combustor inlet static pressure P s31实 and the inlet static temperature T s31实 are calculated; The cold total pressure loss of the combustion chamber is proportional to the square of the inlet converted flow rate, and the combustion chamber inlet converted flow rate Wac at the design point is 3des and the cold total pressure recovery coefficient σ of the combustion chamber 3des The actual total pressure recovery coefficient σ3 of the combustion chamber is calculated according to the following formula: ； Then, based on the actual total pressure P at the combustion chamber inlet... 31实 The actual cold-state outlet total pressure P is calculated using the combustion chamber actual cold-state total pressure recovery coefficient σ3. 32实 The actual Mach number at the combustion chamber inlet is Ma 31实 Actual total import pressure P 31实 The actual interpolated combustion efficiency η of the residual gas coefficient α on the combustion chamber characteristic diagram b实 According to the actual interpolated combustion efficiency η b实 The residual gas coefficient α is used to determine the actual combustion chamber outlet temperature T. 42实 ; According to the flow conservation equation, the combustion chamber inlet area A 31 Actual combustion chamber inlet flow rate Wa 31实 Actual total pressure at the combustion chamber inlet P 31实 Actual total temperature at the combustion chamber inlet T 31实 Calculate the actual combustion chamber inlet airflow velocity V 31实 ; by the combustion chamber outlet area A 32 Actual combustion chamber outlet flow rate Wg 32实 Actual total pressure at the combustion chamber outlet, P 32实 Actual total combustion chamber outlet temperature T 32实 Calculate the actual combustion chamber outlet airflow velocity V 32实 ; by the combustion chamber inlet area A 31 Actual combustion chamber inlet flow rate Wa 31实 Actual total pressure at the combustion chamber inlet P 31实 Actual total temperature at the combustion chamber inlet T 31实 Calculate the actual combustion chamber inlet static pressure Ps 31实 ; According to the momentum conservation calculation of thermal state loss, the known quantity is substituted into the formula to obtain: ， In the formula, the left side of the equal sign is the impulse considering the cold loss, and the right side is the impulse after heating. The actual total pressure P of the combustion chamber outlet is taken as the residual error of the momentum conservation formula 32实 ; That is, the obtained combustion chamber off-design parameters include: actual combustion chamber outlet total temperature T 32实 , actual combustion chamber outlet total pressure P 32实 , and combustion chamber actual outlet flow rate Wg 32实 ; Step 24, the step of calculating the non-design point of the tail nozzle: The tail nozzle inlet cross section parameters are equal to the combustion chamber outlet cross section parameters, the actual tail nozzle inlet total temperature T 41实 , the actual tail nozzle inlet total pressure P 41实 , and the actual tail nozzle inlet flow rate Wg 41实 , and the design point tail nozzle throat area A 43 , the tail nozzle outlet cross section area A 42 ; The throat of the tail nozzle is calculated according to the critical state, and the known parameters are the actual total temperature T of the throat of the tail nozzle 43实 , the actual flow W of the throat of the tail nozzle g43实 , and the throat area A 43 . The flow continuity equation is obtained by substituting the known parameters into the flow continuity equation: ， Calculate the total pressure P8 required by the actual tail nozzle, Finally, the exit velocity V9 of the nozzle is calculated from the actual nozzle exit flow rate W g42实 , the nozzle exit total temperature T 42实 , the nozzle exit area A 42 , and the actual nozzle exit static pressure P s42 , in combination with the actual nozzle exit flow rate W g42实 , the nozzle exit area A 42 , and the actual nozzle exit static pressure P s42 , and the inlet duct inlet flow rate Wa, the inlet duct inlet velocity Vo, as follows: ， Thus, the engine thrust at the non-design point is calculated, i.e. the actual engine thrust Fn 实 ; That is, the off-design parameters of the nozzle are obtained as the actual engine thrust Fn 实 and the total pressure P8 required by the actual nozzle.

4. The method of claim 3, wherein, The calculation of the total pressure P8 required by the actual tail nozzle in said step 24 further comprises the following steps: The nozzle exit has two states, one is subsonic compression to exit, the corresponding static pressure is denoted as P s1 , and the other is supersonic expansion to exit and then experiencing a normal shock at exit, the corresponding static pressure is denoted as P s2 . Actual atmospheric pressure P s0实 Greater than P s1 The entire nozzle is in subsonic state, and the exit static pressure is equal to atmospheric pressure. According to the static pressure P s0 And the area A9, the total pressure P8 required by the tail nozzle is calculated, and the throat Mach number Ma8 is recalculated; Actual atmospheric pressure P s0实 At P s1 and P s2 , there is a positive shock between the throat and the exit, the nozzle exit is subsonic, the exit static pressure equals atmospheric pressure, then the actual nozzle exit total pressure P s0 is calculated from the static pressure P 43 and the nozzle exit area A 42实 , while the nozzle throat total pressure is still P8; Actual atmospheric pressure P s0实 Less than P s2 The nozzle exit is supersonic, and the total pressure P8 and exit area A 43 Actual nozzle exit static pressure P s9 is calculated.

5. The method of claim 3, wherein, The entire non-steady state process described in Step 4 is calculated at Time = 0, without accounting for combustion chamber volume effects, corresponding to the value of the total temperature T4 at the outlet of the combustion chamber with volume taken into account in Step 23, in place of T 32实 , without accounting for volume, in place of Wg g4 32实 .​ 6. The method of claim 1, wherein, The entire non-steady state process described in step four, at Time > 0, taking into account the combustion chamber volume effect, corresponds to the actual combustion chamber outlet total temperature T 32实 whose value is equal to the combustion chamber outlet total temperature T4 taking into account the volume, the combustion chamber actual outlet flow rate Wg 32实 whose value is equal to the flow rate W g4 taking into account the volume.

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