CN116562194B - Ramjet rotary detonation engine thrust evaluation method and system - Google Patents

Abstract

The invention provides a ramjet rotary detonation engine thrust evaluation method and system, which relates to the technical field of rotary detonation engines. When the exhaust nozzle is in a critical state, a plurality of temperatures distributed along the circumferential direction of the combustion chamber are obtained within a preset time. test value, and calculate the static temperature at the combustion chamber outlet. Calculate the Mach number at the combustion chamber outlet based on the combustion chamber outlet area and the nozzle throat area, and calculate the total temperature at the combustion chamber outlet based on the relationship between total temperature and static temperature. According to the temperature Calculate the combustion efficiency of the combustion chamber using the liter method, and iteratively solve the gas physical parameters in the combustion chamber, and solve the aerodynamic thermal parameters of the engine air inlet, the total pressure of the combustion chamber outlet and the aerodynamic thermal parameters of the exhaust nozzle outlet. According to the engine air inlet and exhaust The aerodynamic parameters at the nozzle outlet are used to calculate engine thrust, which is more accurate than numerical simulation to evaluate the thrust of the entire engine. Compared with the free jet test method, it is simpler and more efficient in thrust evaluation.

Description

Ramjet rotary detonation engine thrust evaluation method and system

Technical field

The present invention relates to the technical field of rotating detonation engines, and in particular to a thrust evaluation method and system for a ramjet rotating detonation engine.

Background technique

In the past, the thrust evaluation methods of ramjet rotary detonation engine are usually divided into numerical simulation method and free jet wind tunnel test method. The use of numerical simulation methods to evaluate the thrust of the entire machine relies on the accuracy of the calculation method. In particular, the rotational detonation wave is difficult to accurately simulate, so the evaluation accuracy of the numerical simulation method is low. The free-jet test method can accurately evaluate the thrust performance of the entire engine, but the free-jet test cycle is long, the test process is complex, and the test cost is high, which results in the inability to quickly iterate the engine design. Since the numerical simulation method and the free-jet wind tunnel test method are independent of each other, there is a lack of technical guidance for integrating the two thrust evaluation methods, and it is difficult to take into account the efficiency and accuracy of thrust evaluation.

Contents of the invention

The purpose of the present invention is to provide a ramjet rotary detonation engine thrust evaluation method and system to alleviate the technical problem that it is difficult to balance efficiency and accuracy in the thrust evaluation of the ramjet rotary detonation engine.

In a first aspect, the present invention provides a thrust evaluation method for a ramjet rotary detonation engine, which includes the following steps: when the exhaust nozzle is in a critical state, obtain multiple temperature test values spaced along the circumference of the combustion chamber within a preset time, And calculate the combustion chamber exit static temperature T _s4 ; obtain the combustion chamber exit area A ₄ and the nozzle throat area A _cr , and calculate the combustion chamber exit Mach based on the combustion chamber exit area A ₄ and the nozzle throat area A _cr Number M _a4 ; according to the formula , calculate the total temperature T _t4 at the combustion chamber outlet; calculate the combustion efficiency of the combustion chamber according to the temperature rise method/> , and iteratively solve the gas physical parameters in the combustion chamber; solve the aerodynamic thermal parameters of the engine air inlet; calculate the total pressure p _t4 at the combustion chamber outlet; solve the aerodynamic thermal parameters of the exhaust nozzle outlet; according to the engine air inlet and exhaust nozzle The aerodynamic and thermal parameters of the outlet are used to solve for the engine thrust F.

In conjunction with the first aspect, the present invention provides a first possible implementation of the first aspect, wherein the step of calculating the combustion chamber outlet static temperature T _s4 includes: according to the formula , calculate the combustion chamber outlet static temperature T _s4 , where u is the number of temperature sensors distributed along the circumferential direction of the combustion chamber,/> is the collection time,/> for/> The test value of the i-th temperature sensor within the time period.

Combined with the first aspect, the present invention provides a second possible implementation of the first aspect, wherein the combustion chamber outlet area A ₄ and the nozzle throat area are obtained , according to the combustion chamber outlet area A ₄ and the nozzle throat area The steps to calculate the combustion chamber outlet Mach number M _a4 include: According to the formula /> , calculate the combustion chamber outlet area A ₄ , where r _4i is the diameter of the inner ring of the combustion chamber outlet, r _4e is the diameter of the outer ring of the combustion chamber outlet; according to the formula/> , calculate the nozzle throat area/> , where r _cri is the diameter of the inner ring of the nozzle throat, r _cre is the diameter of the outer ring of the nozzle throat; according to the formula/> , solve for the combustion chamber exit Mach number M _a4 , where, /> is the initial value of the gas specific heat ratio at the combustion chamber outlet.

Combined with the first aspect, the present invention provides a third possible implementation of the first aspect, wherein the step of solving the aerodynamic and thermal parameters of the engine air inlet includes: obtaining the flight altitude H and the flight Mach number M _a0 , according to The obtained flight altitude H and the flight Mach number M _a0 are calculated to obtain the engine air inlet air flow static temperature T _s0 , the engine air inlet static pressure p _s0 and the engine air inlet air flow density ρ ₀ ; according to the formula , calculate the flight speed v ₀ where, /> is the specific heat of air, R _a is the gas constant of air; according to the formula/> , calculate the air flow rate m _a of the engine, where, is the intake capture area of the engine,/> is the engine inlet flow coefficient; according to the formula , calculate the total airflow temperature T _t0 at the inlet of the inlet.

In combination with the third possible implementation of the first aspect, the present invention provides a fourth possible implementation of the first aspect, wherein the calculation obtains the static temperature of the engine air inlet airflow , engine air inlet static pressure/> and engine air inlet airflow density/> The steps include: compare and judge the range of flight altitude H; when 0<H<11000m, ,/> ;When 11000m≤H<24000m, ,/> ;According to the formula/> , calculate the airflow density ρ ₀ at the engine air inlet, where g is the acceleration of gravity,/> is the atmospheric pressure at sea level, and e is the natural logarithm.

Combined with the first aspect, the present invention provides a fifth possible implementation of the first aspect, wherein the combustion chamber combustion efficiency is calculated according to the temperature rise method. The steps include: According to the formula/> , calculate combustion efficiency/> , where,/> is the constant pressure specific heat value of the combustion chamber outlet gas,/> is the lower calorific value of fuel,/> is the total temperature at the combustion chamber inlet,/> ,/> is the constant pressure specific heat value of the imported air,/> is the specific heat value of fuel.

In conjunction with the first aspect, the present invention provides a sixth possible implementation of the first aspect, wherein the calculation of the total pressure at the combustion chamber outlet The steps include: According to the formula/> , calculate the fuel flow rate involved in combustion/> , where,/> is the fuel flow into the engine;

According to the formula , calculate the oxygen flow rate involved in combustion/> , where L ₀ is the theoretical amount of air required to completely burn 1 kilogram of fuel; according to the formula/> , calculate the flow rate of carbon dioxide in the combustion products , where,/> is the molecular mass of carbon dioxide,/> is the molecular mass of fuel; according to the formula , calculate the water vapor flow rate in the combustion products, where,/> is the molecular mass of water vapor; according to the formula/> , calculate the nitrogen flow rate in the air that does not participate in combustion/> , where,/> is the mass fraction of nitrogen; according to the formula/> , calculate the oxygen flow rate in the air that does not participate in combustion/> , where,/> is the mass fraction of oxygen; according to the formula/> , calculate the mass proportion Y _j of each component gas in the combustion chamber, where j component includes nitrogen, oxygen, carbon dioxide, water vapor and kerosene vapor, n=5; according to the formula , calculate the average gas constant R _ave of the gas mass in the combustion chamber, where R _j is the gas constant of component j; according to the formula/> , calculate the constant-pressure specific heat C _{p,ave of} the average gas mass at the combustion chamber outlet, where C _Pj is the constant-pressure specific heat of j component at the combustion chamber outlet, /> , a, b, c, d, e and f are polynomial calculation parameters respectively; according to the formula/> , calculate the specific heat at constant volume C _v,ave ; according to the formula , calculate the specific heat ratio/> ;Compare specific heat ratio calculations/> Initial value of specific heat ratio to combustion chamber outlet gas Whether the difference is less than or equal to the preset difference ε, if not, the specific heat ratio will be calculated/> Assigned to the initial value of the gas specific heat ratio at the combustion chamber outlet/> , iteratively calculate the specific heat ratio/> , until the calculated value of the specific heat ratio/> and the initial value of the specific heat ratio of the combustion chamber outlet gas/> The difference is less than or equal to the preset difference ε, and the final total combustion chamber outlet temperature T _t4 is obtained; according to the formula/> , calculate the total pressure at the combustion chamber outlet/> , where,/> , ,/> is the velocity coefficient at the exhaust nozzle throat, under the critical or supercritical state of the exhaust nozzle throat/> .

Combined with the first aspect, the present invention provides a seventh possible implementation of the first aspect, wherein the step of solving the aerodynamic thermal parameters of the exhaust nozzle outlet includes: according to the formula , calculate the exhaust nozzle exit Mach number/> ;According to the formula , calculate the nozzle outlet static pressure/> ,/> is the total pressure at the nozzle outlet, assuming absolute isentropic flow/> ;According to the formula/> , calculate the combustion chamber outlet static temperature/> , where,/> is the total temperature of the nozzle; according to the formula/> , calculate the exhaust nozzle exit velocity/> .

Combined with the first aspect, the present invention provides an eighth possible implementation of the first aspect, wherein the step of solving the engine thrust F according to the aerodynamic parameters of the engine air inlet and the exhaust nozzle outlet includes: formula , calculate the engine thrust F, where,/> is the nozzle impulse,/> ;/> is the engine air inlet impulse,/> .

In a second aspect, the present invention provides a ramjet rotary detonation engine thrust evaluation system, including: a processor and a plurality of temperature sensors. The processor is used to execute a program corresponding to the ramjet rotary detonation engine thrust evaluation method recorded in the first aspect. ; A plurality of the temperature sensors are respectively installed in the flame tube of the ramjet rotary detonation engine, and are arranged at intervals along the circumference of the combustion chamber; a plurality of the temperature sensors are respectively connected to the processor.

The embodiments of the present invention bring the following beneficial effects: when the exhaust nozzle is in a critical state, multiple temperature test values distributed along the circumference of the combustion chamber within a preset time are obtained, and the static temperature at the combustion chamber outlet is calculated. The combustion chamber outlet area and the nozzle throat area are used to calculate the combustion chamber outlet Mach number, and the total temperature at the combustion chamber outlet is calculated according to the relationship between total temperature and static temperature. The combustion chamber combustion efficiency is calculated according to the temperature rise method, and iteratively solves the problem in the combustion chamber. Gas physical property parameters, solve the aerodynamic thermal parameters of the engine air inlet, the total pressure of the combustion chamber outlet and the aerodynamic thermal parameters of the exhaust nozzle outlet, and solve the engine thrust based on the aerodynamic thermal parameters of the engine air inlet and exhaust nozzle outlet, correspondingly Compared with numerical simulation, it is more accurate to evaluate the thrust of the whole machine. Compared with the free jet test method, it is simpler and more efficient in thrust evaluation.

In order to make the above-mentioned objects, features and advantages of the present invention more obvious and understandable, preferred embodiments are given below and described in detail with reference to the accompanying drawings.

Description of drawings

In order to more clearly explain the specific embodiments of the present invention or the technical solutions in related technologies, the drawings that need to be used in the description of the specific embodiments or related technologies will be briefly introduced below. Obviously, the drawings in the following description are: For some embodiments of the present invention, those of ordinary skill in the art can also obtain other drawings based on these drawings without exerting creative efforts.

Figure 1 is a schematic flow chart of a ramjet rotary detonation engine thrust evaluation method provided by an embodiment of the present invention;

Figure 2 is a schematic diagram of a ramjet rotary detonation engine;

Figure 3 is a cross-sectional view of a ramjet rotary detonation engine.

Icon: 1-injection section; 2-diffusion section; 3-flame barrel; 4-exhaust nozzle; 5-nozzle; 6-hot jet detonation device; 7-temperature sensor.

Detailed ways

The technical solution of the present invention will be clearly and completely described below with reference to the accompanying drawings. Obviously, the described embodiments are some, not all, of the embodiments of the present invention. Based on the embodiments of the present invention, all other embodiments obtained by those of ordinary skill in the art without creative efforts fall within the scope of protection of the present invention.

In the description of the present invention, it should be noted that the terms "center", "upper", "lower", "left", "right", "vertical", "horizontal", "inner", "outer", etc. The indicated orientation or positional relationship is based on the orientation or positional relationship shown in the drawings. It is only for the convenience of describing the present invention and simplifying the description. It does not indicate or imply that the device or element referred to must have a specific orientation or a specific orientation. construction and operation, and therefore should not be construed as limitations of the invention. Furthermore, the terms “first”, “second” and “third” are used for descriptive purposes only and are not to be construed as indicating or implying relative importance. The physical quantities in the formula, if not separately marked, should be understood as the basic quantities in the basic units of the International System of Units, or the derived quantities derived from the basic quantities through mathematical operations such as multiplication, division, differentiation or integration.

In the description of the present invention, it should be noted that, unless otherwise clearly stated and limited, the terms "installation", "connection" and "connection" should be understood in a broad sense. For example, it can be a fixed connection or a detachable connection. Connection, or integral connection; it can be a mechanical connection or an electrical connection; it can be a direct connection or an indirect connection through an intermediate medium; it can be an internal connection between two components. For those of ordinary skill in the art, the specific meanings of the above terms in the present invention can be understood on a case-by-case basis.

As shown in Figure 1, the thrust evaluation method of a ramjet rotary detonation engine provided by an embodiment of the present invention includes the following steps:

When the exhaust nozzle is in a critical state, obtain multiple temperature test values distributed along the circumference of the combustion chamber within a preset time, and calculate the combustion chamber outlet static temperature T _s4 ;

Obtain the combustion chamber exit area A ₄ and the nozzle throat area A _cr , and calculate the combustion chamber exit Mach number M _a4 based on the combustion chamber exit area A ₄ and the nozzle throat area A _cr ;

According to the formula , calculate the total temperature T _t4 at the combustion chamber outlet;

Calculate combustion chamber combustion efficiency based on temperature rise method , and iteratively solve the gas physical parameters in the combustion chamber;

Solve the aerodynamic and thermal parameters of the engine air inlet;

Calculate the total pressure p _t4 at the combustion chamber outlet;

Solve the aerodynamic and thermal parameters of the exhaust nozzle outlet;

According to the aerodynamic and thermal parameters of the engine inlet and exhaust nozzle outlet, the engine thrust F is solved.

The thrust evaluation method of the ramjet rotary detonation engine recorded in this embodiment combines the combustion chamber model test data and the aerodynamics mathematical model analysis method. Through component-level tests, the thrust of the entire ramjet rotary detonation engine can be evaluated. The evaluation method Accurate, simple and efficient assessment process. Compared with numerical simulation, it is more accurate to evaluate the thrust of the whole engine. Compared with the free jet test method, it is simpler and more efficient in thrust evaluation. This makes the engine thrust evaluation both efficient and accurate.

In the embodiment of the present invention, the steps of calculating the combustion chamber outlet static temperature T _s4 include: according to the formula , calculate the combustion chamber outlet static temperature T _s4 , where u is the number of temperature sensors distributed along the circumferential direction of the combustion chamber,/> is the collection time,/> for/> The test value of the i-th temperature sensor within the time period. The accuracy of the average static temperature in the circumferential direction of the combustion chamber outlet can be improved by increasing the number of temperature sensors. u can take a value of 4, 5, 6 or more. Multiple temperature sensors are evenly spaced along the circumferential direction of the combustion chamber. In time and Solving the average static pressure in space solves the problems of uneven distribution of static temperature at the combustion chamber outlet and time non-steadiness of static temperature measurement.

Further, obtain the combustion chamber outlet area A ₄ and the nozzle throat area , based on the combustion chamber outlet area A ₄ and the nozzle throat area/> The steps to calculate the combustion chamber exit Mach number M _a4 include:

According to the formula , calculate the combustion chamber outlet area A ₄ , where r _4i is the diameter of the inner ring of the combustion chamber outlet, r _4e is the diameter of the outer ring of the combustion chamber outlet;

According to the formula , calculate the nozzle throat area/> , where r _cri is the diameter of the inner ring of the nozzle throat, r _cre is the diameter of the outer ring of the nozzle throat;

According to the formula , solve for the combustion chamber exit Mach number M _a4 , where, /> is the initial value of the specific heat ratio of the combustion chamber outlet gas, and the initial value of the combustion chamber outlet gas specific heat ratio/> It can be preset or input in advance, r _4i , r _4e , r _cri and r _cre can be preset according to the pressure rotation detonation engine or detected in the test.

Further, the steps to solve the aerodynamic and thermal parameters of the engine air inlet include:

Obtain the flight altitude H and flight Mach number M _a0 , and calculate the engine air inlet air flow static temperature T _s0 , engine air inlet static pressure p _s0 and engine air inlet air flow density ρ according to the flight altitude H and flight Mach number M _a0 . ₀ ;

According to the formula , calculate the flight speed v ₀ where, /> is the specific heat of air, R _a is the gas constant of air;

According to the formula , calculate the engine air flow m _a , where,/> is the intake capture area of the engine,/> is the engine inlet flow coefficient;

According to the formula , calculate the total airflow temperature T _t0 at the inlet of the inlet.

Further, calculate and obtain the static temperature of the engine air inlet airflow. , engine air inlet static pressure/> and engine air inlet airflow density/> The steps include:

Compare and determine the range of flight altitude H;

When 0<H<11000m, ,/> ;

When 11000m≤H<24000m, ,/> ;

According to the formula , calculate the airflow density ρ ₀ at the engine air inlet, where g is the acceleration of gravity,/> is the atmospheric pressure at sea level, and e is the natural logarithm.

Further, calculate the combustion efficiency of the combustion chamber according to the temperature rise method. The steps include:

According to the formula , calculate combustion efficiency/> , where,/> is the constant pressure specific heat value of the combustion chamber outlet gas,/> is the lower calorific value of fuel,/> is the total temperature at the combustion chamber inlet,/> ,/> is the constant pressure specific heat value of the imported air,/> is the specific heat value of fuel.

Further, calculate the total pressure at the combustion chamber outlet The steps include:

According to the formula , calculate the fuel flow rate involved in combustion/> , where,/> is the fuel flow into the engine;

According to the formula , calculate the oxygen flow rate involved in combustion/> , where L ₀ is the theoretical amount of air required to completely burn 1 kilogram of fuel;

According to the formula , calculate the flow rate of carbon dioxide in the combustion products/> , where,/> is the molecular mass of carbon dioxide,/> is the molecular weight of fuel;

According to the formula , calculate the water vapor flow rate in the combustion products, where,/> is the molecular mass of water vapor;

According to the formula , calculate the nitrogen flow rate in the air that does not participate in combustion/> ,in, is the mass fraction of nitrogen;

According to the formula , calculate the oxygen flow rate in the air that does not participate in combustion/> ,in, is the mass fraction of oxygen;

According to the formula , calculate the mass proportion Y _j of each component gas in the combustion chamber, where j component includes nitrogen, oxygen, carbon dioxide, water vapor and kerosene vapor, n=5;

According to the formula , calculate the average gas constant R _ave of the gas mass in the combustion chamber, where R _j is the gas constant of component j;

According to the formula , calculate the constant-pressure specific heat C _{p,ave of} the average gas mass at the combustion chamber outlet, where C _Pj is the constant-pressure specific heat of j component at the combustion chamber outlet, /> , a, b, c, d, e and f are polynomial calculation parameters respectively;

According to the formula , calculate the specific heat at constant volume C _v,ave ;

According to the formula , calculate the specific heat ratio/> ;

Compare specific heat ratio calculations Initial value of specific heat ratio to combustion chamber outlet gas/> Whether the difference is less than or equal to the preset difference ε, if not, the specific heat ratio will be calculated/> Assigned to the initial value of the gas specific heat ratio at the combustion chamber outlet/> , iteratively calculate the specific heat ratio/> , until the specific heat ratio is calculated/> Initial value of specific heat ratio to combustion chamber outlet gas/> The difference is less than or equal to the preset difference ε, and the final total combustion chamber outlet temperature T _t4 is obtained;

According to the formula , calculate the total pressure at the combustion chamber outlet/> , where,/> , ,/> is the velocity coefficient at the exhaust nozzle throat, under the critical or supercritical state of the exhaust nozzle throat/> .

Further, the steps to solve the aerodynamic parameters of the exhaust nozzle outlet include:

According to the formula , calculate the exhaust nozzle exit Mach number/> ;

According to the formula , calculate the nozzle outlet static pressure/> ,/> is the total pressure at the nozzle outlet, assuming absolute isentropic flow/> ;

According to the formula , calculate the combustion chamber outlet static temperature/> , where,/> is the total temperature of the nozzle;

According to the formula , calculate the exhaust nozzle exit velocity/> .

Further, based on the aerodynamic parameters of the engine air inlet and exhaust nozzle outlet, the steps to solve the engine thrust F include:

According to the formula , calculate the engine thrust F, where,/> is the nozzle impulse,/> ; is the engine air inlet impulse,/> . From this it can be deduced that the calculated engine thrust .

Using the above ramjet rotary detonation engine thrust evaluation method, although the specific heat and specific heat ratio of the gas mixture in the combustion chamber change greatly with temperature and composition, in the ramjet rotary detonation engine thrust evaluation method, the static temperature at the combustion chamber outlet is The combustion chamber flow rate is used as a constraint, and the physical property parameters such as the specific heat ratio of the mixed gas and the aerodynamic thermodynamic parameters such as the total pressure of the mixed gas are iteratively solved to obtain the engine thrust. On the one hand, the combustion chamber model is simplified, the test process is simple, the test cycle is short, and the test cost is low; on the other hand, the total pressure and total temperature at the combustion chamber outlet are calculated accurately, which improves the accuracy of thrust evaluation.

As shown in Figures 2 and 3, the ramjet rotary detonation engine thrust evaluation system provided by the embodiment of the present invention includes: a processor and multiple temperature sensors 7. Each step of the ramjet rotary detonation engine thrust evaluation method can be integrated, programmed according to the flow chart shown in Figure 1, and the program can be implanted into the memory and the processor can program the program. The rotary detonation engine to which this system is applicable has an injection section 1, a diffusion section 2, a flame tube 3 and an exhaust nozzle 4 arranged in sequence. The tubes 4 are all composed of an inner tube and an outer tube. In the injection section 1, the diameter of the inner tube and the outer tube is constant along the axial direction. In the expansion section 2, the diameter of the inner tube gradually decreases and the diameter of the outer tube gradually increases along the air flow direction. In the flame The diameter of the outer tube and the inner tube at barrel 3 remains constant along the direction of air flow. At exhaust nozzle 4, the diameter of the inner tube first increases and then decreases along the direction of air flow, while the diameter of the outer tube first decreases and then increases. Air is taken in from the axial end of the injection section 1 away from the diffuser section 2, and fuel is injected into it from a plurality of nozzles 5 installed in the injection section 1. The plurality of nozzles 5 are arranged at intervals along the circumference of the injection section 1 . The oil-gas mixture enters the flame tube 3, and the annular chamber inside the flame tube 3 is used as the combustion chamber. The hot jet detonation device 6 installed in the flame tube 3 close to the expansion section 2 is ignited and detonated, and a rotation can be formed in the flame tube 3. Detonation wave. The exhaust gas generated by the detonation of the flame tube 3 flows into the exhaust nozzle 4 and is ejected through the end of the exhaust nozzle 4. A plurality of temperature sensors 7 are respectively installed on the flame tube 3 of the ramjet rotary detonation engine and are arranged at intervals along the circumference of the combustion chamber. The plurality of temperature sensors 7 are respectively connected to the processor. Interactive devices can also be added and connected to the processor.

Using the above-mentioned ramjet rotary detonation engine thrust evaluation method and system has the following beneficial effects:

(1) Combining combustion chamber model test data and aerodynamic thermodynamic mathematical model analysis methods, through component-level tests, the thrust of the entire ramjet rotary detonation engine is evaluated. The evaluation method is accurate and the evaluation process is simple. Compared with the method of evaluating the thrust of the whole machine through numerical simulation, the present invention conducts evaluation based on the test data of rotating detonation combustion, which has high accuracy. Compared with the free jet test, the combustion chamber model of the present invention is simple, the test process is simple, the test period is short, and the test cost is low.

(2) Multiple temperature detectors are evenly distributed around the circumference of the combustion chamber to collect the static temperature at the combustion chamber outlet, which solves the problem of spatial non-uniformity in the static temperature at the combustion chamber outlet. In the thrust evaluation method, the static temperature of multiple periods collected by the temperature sensor is time averaged to obtain the static temperature of the combustion chamber outlet, which solves the problem of time non-steadiness in static temperature measurement.

(3) Using the combustion chamber outlet static temperature and combustion chamber flow rate as constraints, iteratively solve the physical property parameters such as the specific heat ratio of the mixed gas and the aerodynamic thermodynamic parameters such as the total pressure of the mixed gas, and then solve the overall engine thrust. The influence of temperature on the physical property parameters specific heat ratio and gas constant is specifically avoided, the total pressure and total temperature at the combustion chamber outlet are calculated more accurately, and the engine thrust evaluation is more accurate.

Finally, it should be noted that the above embodiments are only used to illustrate the technical solution of the present invention, but not to limit it. Although the present invention has been described in detail with reference to the foregoing embodiments, those of ordinary skill in the art should understand that: The technical solutions described in the foregoing embodiments can still be modified, or some or all of the technical features can be equivalently replaced; and these modifications or substitutions do not deviate from the essence of the corresponding technical solutions from the technical solutions of the embodiments of the present invention. scope.

Claims (7)

1. A ramjet rotary detonation engine thrust evaluation method, characterized by including the following steps: When the exhaust nozzle is in a critical state, obtain multiple temperature test values distributed along the circumference of the combustion chamber within a preset time, and calculate the combustion chamber outlet static temperature T _s4 ; Obtain the combustion chamber outlet area A ₄ and the nozzle throat area A _cr , and calculate the combustion chamber outlet Mach number M _a4 based on the combustion chamber outlet area A ₄ and the nozzle throat area A _cr ; According to the formula , calculate the total temperature T _t4 at the combustion chamber outlet; Calculate combustion chamber combustion efficiency based on temperature rise method , and iteratively solve the gas physical parameters in the combustion chamber; Solve the aerodynamic and thermal parameters of the engine air inlet; the step of solving the aerodynamic and thermal parameters of the engine air inlet includes: obtaining the flight altitude H and the flight Mach number M _a0 , and obtaining the flight altitude H and the flight Mach number M according to _a0 is calculated to obtain the engine air inlet air flow static temperature T _s0 , engine air inlet static pressure p _s0 and engine air inlet air flow density ρ ₀ ; according to the formula , calculate the flight speed v ₀ where, /> is the specific heat of air, R _a is the air gas constant; according to the formula , calculate the engine air flow m _a , where,/> is the intake capture area of the engine,/> is the engine inlet flow coefficient; according to the formula/> , calculate the total airflow temperature T _t0 at the inlet of the inlet; Calculate the total pressure p _t4 at the combustion chamber outlet; the steps of calculating the total pressure p _t4 at the combustion chamber outlet include: according to the formula , calculate the fuel flow rate involved in combustion/> , where,/> is the fuel flow entering the engine; according to the formula , calculate the oxygen flow rate involved in combustion/> , where L ₀ is the theoretical amount of air required to completely burn 1 kilogram of fuel; according to the formula/> , calculate the flow rate of carbon dioxide in the combustion products/> , where,/> is the molecular mass of carbon dioxide,/> is the molecular mass of fuel; according to the formula/> , calculate the water vapor flow rate in the combustion products, where,/> is the molecular mass of water vapor; according to the formula/> , calculate the nitrogen flow rate in the air that does not participate in combustion/> , where,/> is the mass fraction of nitrogen; according to the formula , calculate the oxygen flow rate in the air that does not participate in combustion/> , where,/> is the mass fraction of oxygen; according to the formula/> , calculate the mass proportion Y _j of each component gas in the combustion chamber, where j component includes nitrogen, oxygen, carbon dioxide, water vapor and kerosene vapor, n=5; according to the formula/> , calculate the average gas constant R _ave of the gas mass in the combustion chamber, where R _j is the gas constant of component j; according to the formula/> , calculate the constant pressure specific heat C _{p,ave of} the average gas mass at the combustion chamber outlet, where C _Pj is the constant pressure specific heat of component j at the combustion chamber outlet, , a, b, c, d, e and f are polynomial calculation parameters respectively; according to the formula , calculate the specific heat at constant volume C _v,ave ; according to the formula/> , calculate the specific heat ratio/> ;Compare specific heat ratio calculations/> Initial value of specific heat ratio to combustion chamber outlet gas/> Whether the difference is less than or equal to the preset difference ε, if not, the specific heat ratio will be calculated/> Assigned to the initial value of the gas specific heat ratio at the combustion chamber outlet/> , iteratively calculate the specific heat ratio/> , until the calculated value of the specific heat ratio/> and the initial value of the specific heat ratio of the combustion chamber outlet gas/> The difference is less than or equal to the preset difference ε, and the final total combustion chamber outlet temperature T _t4 is obtained; according to the formula/> , calculate the total pressure at the combustion chamber outlet/> , where,/> ,/> ,/> is the velocity coefficient at the exhaust nozzle throat, under the critical or supercritical state of the exhaust nozzle throat/> ; Solve the aerodynamic thermal parameters of the exhaust nozzle outlet; the steps of solving the aerodynamic thermal parameters of the exhaust nozzle outlet include: according to the formula , calculate the exhaust nozzle exit Mach number/> ;According to the formula/> , calculate the nozzle outlet static pressure/> ,/> is the total pressure at the nozzle outlet, assuming absolute isentropic flow/> ;According to the formula/> , calculate the combustion chamber outlet static temperature/> , where,/> is the total temperature of the nozzle; according to the formula/> , calculate the exhaust nozzle exit velocity/> ; According to the aerodynamic and thermal parameters of the engine inlet and exhaust nozzle outlet, the engine thrust F is solved. 2. Ramjet rotary detonation engine thrust evaluation method according to claim 1, characterized in that the step of calculating the combustion chamber outlet static temperature T _s4 includes: According to the formula , calculate the combustion chamber outlet static temperature T _s4 , where u is the number of temperature sensors distributed along the circumferential direction of the combustion chamber,/> is the collection time,/> for/> The test value of the i-th temperature sensor within the time period. 3. Ramjet rotary detonation engine thrust evaluation method according to claim 1, characterized in that said obtaining combustion chamber outlet area _A4 and nozzle throat area , according to the combustion chamber outlet area A ₄ and the nozzle throat area/> The steps to calculate the combustion chamber exit Mach number M _a4 include: According to the formula , calculate the combustion chamber outlet area A ₄ , where r _4i is the diameter of the inner ring of the combustion chamber outlet, r _4e is the diameter of the outer ring of the combustion chamber outlet; According to the formula , calculate the nozzle throat area/> , where r _cri is the diameter of the inner ring of the nozzle throat, r _cre is the diameter of the outer ring of the nozzle throat; According to the formula , solve for the combustion chamber exit Mach number M _a4 , where, /> is the initial value of the gas specific heat ratio at the combustion chamber outlet. 4. The ramjet rotary detonation engine thrust evaluation method according to claim 1, characterized in that the calculation obtains the engine air inlet air flow static temperature T _s0 , the engine air inlet static pressure p _s0 and the engine air inlet air flow. The steps for density ρ ₀ include: Compare and determine the range of flight altitude H; When 0<H<11000m, ,/> ; When 11000m≤H<24000m, ,/> ; According to the formula , calculate the airflow density ρ ₀ at the engine air inlet, where g is the acceleration of gravity,/> is the atmospheric pressure at sea level, and e is the natural logarithm. 5. The ramjet rotary detonation engine thrust evaluation method according to claim 1, characterized in that the combustion chamber combustion efficiency is calculated according to the temperature rise method. The steps include: According to the formula , calculate combustion efficiency/> , where,/> is the constant pressure specific heat value of the combustion chamber outlet gas,/> is the lower calorific value of fuel,/> is the total temperature at the combustion chamber inlet,/> ,/> is the constant pressure specific heat value of the imported air,/> is the specific heat value of fuel. 6. The ramjet rotary detonation engine thrust evaluation method according to claim 1, wherein the step of solving the engine thrust F according to the aerodynamic parameters of the engine air inlet and the exhaust nozzle outlet includes: According to the formula , calculate the engine thrust F, where,/> is the nozzle impulse,/> ;/> is the engine air inlet impulse,/> . 7. A ramjet rotary detonation engine thrust evaluation system, characterized by including: A processor, the processor being configured to execute a program corresponding to the ramjet rotary detonation engine thrust evaluation method according to any one of claims 1-6; A plurality of temperature sensors (7). The plurality of temperature sensors (7) are respectively installed on the flame tube (3) of the ramjet rotary detonation engine and are arranged at intervals along the circumference of the combustion chamber; Wherein, a plurality of the temperature sensors (7) are respectively connected to the processor.