Thermal flow testing device and method for flight environment simulation of attitude control engine

Abstract

The application provides a heat flow test device and method for simulating the heat flow of a flight environment of a attitude control engine, and mainly solves the technical problem of large measurement error of the existing heat flow measuring device. The heat flow test device comprises a fixing seat, a displacement adjusting seat and an angle adjusting seat arranged below the fixing seat, and a heat insulation layer, a temperature sensor and a sensitive layer arranged above the fixing seat in sequence. The test method is to obtain the thermal effect data of different scales of the attitude control engine under the ignition state by inverting the real plume thermal effect data of the attitude control engine in orbit, combining the radiant heat flow of the quartz lamp array and the external radiant heat flow, realizing the measurement of the thermal flow gradient of the heat environment at the ignition moment of the engine, ensuring the continuity and reliability of the heat flow loading in the test process, and providing a guarantee for verifying the working performance of the main engine of a certain spacecraft under the simultaneous action of the working state and the external heat flow.

Description

TECHNICAL FIELD

[0001] The present application relates to a liquid rocket engine environment simulation test method, in particular to a flight environment simulation heat flow testing device and method for attitude control engine. BACKGROUND

[0002] The attitude control power system is mainly used for attitude control of a hypersonic vehicle, and can also be used for speed correction, orbit adjustment, position keeping, etc. of the hypersonic vehicle, and is an indispensable power device for orbit insertion, re-entry, final correction, etc. of the hypersonic vehicle, and plays an important role in ensuring the hitting accuracy of the weapon type. The long-time flight of the attitude control power system in the atmosphere will cause serious aerodynamic heating, and the thermal control environment of the re-entry atmosphere is more complex, and the heat force is more obvious. The flow and the combustion caused vibration and impact form a complex force-heat coupling environment, and this environment is the most important and critical flight phase of the boost-glide vehicle, and is also the most severe flight phase. The environment can have a great impact on the structure and material of the hypersonic vehicle, and is particularly important for the reliable work of the vehicle and the striking precision.

[0003] At present, the vacuum environment simulation test of the engine is generally carried out under the condition that the engine is not ignited for static heat test, so as to realize the simulation of the vacuum thermal environment, and the research on the vacuum force-heat coupling simulation environment of the engine ignition condition (the engine ignition time is more than 200s, and the thrust is 10-2000N) is less. Under the long-time hypersonic flight environment, the structure of the engine may cause more obvious coupling phenomenon due to the deformation caused by the long-time ignition of the gas combustion and the friction heating, and the cumulative heating of the structure is easy to affect the thermal protection of the vehicle, and even affect the heat transfer characteristics and reliability of the structure inside including the cabin. At the same time, due to the non-uniformity and instability of the engine combustion, and the continuous shock wave generated by the long-time ignition of the engine, the reverse impact on the vacuum force-heat environment vehicle is formed, and then the structural stability and the matching between the system components are affected. Compared with the static test and the short-time engine ignition verification test, the vacuum force-heat coupling simulation environment test under the engine ignition condition is more critical for the accurate evaluation of the flight environment state, and therefore it is necessary to carry out the long-range ignition dynamic heat test of the attitude control engine, so as to improve the accurate evaluation ability of the flight environment state, and adapt to the real flight verification demand of the long-time hypersonic vehicle.

[0004] The existing technology generally adopts the transient heat flow density application method for the research and development of the vacuum force-heat coupling simulation environment under the engine ignition condition, but it has the following problems:

[0005] 1) In the simulation of the coupled thermal environment of the attitude control power system, quartz lamp radiation is generally used to simulate rapid aerodynamic heating. However, due to the instantaneous ignition of the engine, a transient thermal environment with coupled thermal environment is formed. The heating loading characteristics of this environment are significantly different from those of high-temperature and normal-temperature environments. In this transient thermal environment with coupled thermal environment, when the actual heat flux is applied, each partition will have uneven heat flow at the control point and the partition edge.

[0006] 2) Steady-state calibration obtains the relationship between the quartz lamp power current and the calorimeter temperature when equilibrium is reached. However, in the instantaneous mechanical and thermal environment of engine ignition, a rapid thermal equilibrium test can form a transient operating stage. However, affected by the engine heat capacity and the simulated vacuum at high altitude, the high-altitude engine jet flow field will expand and reverse, thereby forming convective heating on the bottom solid surface; at the same time, the high-temperature jet and nozzle solid wall will heat the engine surroundings or payload through thermal radiation. The transient heat flux density reaching the surface is constantly changing, and it takes a long time to reach the required heat flux density. This heat flux non-uniformity causes the heating process of the specimen surface in the heating zone to differ from the actual flight conditions, which may lead to "over-testing" or "under-testing" phenomena, thereby introducing errors into the performance verification process.

[0007] In summary, the existing mechanical-thermal coupling transient thermal environment test formed by engine ignition uses a general transient heat flux density application method, which has certain errors. It is necessary to measure and analyze the transient mechanical-thermal environment heat flux loading gradient. Summary of the Invention

[0008] The purpose of the present invention is to solve the technical problems such as large measurement errors of existing heat flow measurement devices, and to provide a heat flow testing device and method for simulating heat flow in the flight environment of an attitude control engine.

[0009] To achieve the above object, the technical solution of the present invention is as follows:

[0010] A thermal flow test device for simulating thermal flow in the flight environment of an attitude control engine is characterized in that: the thermal flow test device is used to be installed on the outside of a nozzle affected by the attitude control engine, and includes a fixing seat, a displacement adjustment seat and an angle adjustment seat respectively arranged below the fixing seat, and a thermal insulation layer, a temperature sensor, and a sensitive layer arranged above the fixing seat;

[0011] The displacement adjustment seat is used to realize the movement of the fixed seat during the measurement process, and the angle adjustment seat is used to realize the installation angle of the fixed seat during the measurement process;

[0012] The thermal insulation layer covers the top of the fixing seat, and the thermal insulation layer includes a plurality of thermal insulation surfaces covered in sequence, which are used to reduce the heat conduction from the sensitive surface to the surrounding area after absorbing the heat flow;

[0013] The temperature sensor is located above the thermal insulation layer; the sensitive layer is wrapped around the temperature sensor and faces the attitude control engine, and is used to absorb the heat flow of the engine and convert it into its own temperature. The temperature sensor then collects this temperature information and outputs it to an external device.

[0014] Furthermore, the thickness of the sensitive layer is 50 to 100 μm, and a high-temperature resistant and anti-oxidation coating is provided above the sensitive layer, and the high-temperature resistant and anti-oxidation coating faces the attitude control engine.

[0015] Furthermore, the sensitive layer is made of stainless steel foil; and the heat insulation layer is formed by laminating double-sided aluminum-plated polyimide films.

[0016] In addition, the present invention provides a method for testing heat flow in a simulated flight environment of an attitude control engine, using the above-mentioned heat flow testing device for simulating heat flow in a flight environment of an attitude control engine, and specifically comprising the following steps:

[0017] Step 1: Use a quartz lamp array to simulate the gas jet process of an attitude control engine in a flight environment, and use a heat flux test device to measure the radiation heat flux Q2 of the quartz lamp array at different heights, different spacings, and different voltages;

[0018] Step 2: Based on the radiation heat flux Q2 measured in step 1, the uniformity distribution of the heat flux loading is obtained, thereby obtaining the sensitivity of the heat flux loading;

[0019] Step 3: Use a heat flow test device to perform heat flow tests on the engine in different working conditions and at different parts in the axial and radial directions to obtain the heat flow q of the engine in different working conditions. t , and the rate of change of axial and radial heat flux gradients of the engine under different operating conditions;

[0020] Step 4: Compare the engine gas heat flux under different operating conditions obtained in Step 3 with the gas heat flux value during flight. If the error between the two is within 1% to 5%, proceed to Step 5; otherwise, correct the heat capacity value of the heat flux testing device and return to Step 1.

[0021] Step 5: Test the heat flux output value of the engine under different working conditions using a heat flux testing device to obtain the reflectivity of the heat flux testing device, thereby obtaining the outward radiation heat flux Q4 of the sensitive layer under different working conditions;

[0022] Step 6: By inverting the actual heat flux data of the attitude control engine on orbit, combined with the radiation heat flux Q2 of the quartz lamp array obtained in step 1 and the outward radiation heat flux Q4 of the sensitive layer obtained in step 5, as well as the heat flux Q1 of the infrared heat flux emitted by the quartz lamp array and absorbed by the sensitive layer, and the conduction heat flux Q3 of the sensitive layer, the heat flux gradient in the thermal environment at the moment of ignition of the attitude control engine is calculated.

[0023] Furthermore, in step 1, the calculation method of the radiation heat flux Q2 of the quartz lamp array at different heights, different spacings and different voltages is:

[0024]

[0025] Where: q s (x, y, z) is the radiation heat flux Q2, which is the radiation heat flux at the coordinate (x, y, z) on the plate, Q0 is the quartz lamp emission power, H is the distance between the center of the quartz lamp and the plate, L is the heating length of the quartz lamp filament; α, β, and γ are all custom values, and is the angle between the heat flow measurement device and the engine axis.

[0026] Furthermore, step 2 is specifically to obtain the uniformity distribution of the heat flux loading based on the radiation heat flux Q2 measured in step 1, and then obtain the non-uniformity of the heat flux loading. The non-uniformity calculation formula is as follows:

[0027]

[0028] Where, δ is the unevenness of the quartz lamp array on the illuminated surface; q max is the maximum heat flux density on the illuminated surface; q min is the minimum heat flux density on the illuminated surface; at the same time, different influencing factors are combined to obtain the induction coefficient of different influencing factors, thereby obtaining the sensitivity of heat flux loading; the different influencing factors include the distribution distance, angle, working environment and vacuum degree of the quartz lamp array; the sensitivity of the heat flux loading Expressed as:

[0029]

[0030] in: That is, it represents the sensitivity of heat flux loading under different influencing factors.

[0031] Furthermore, in step 3, the calculation formula of the gas heat flow of the engine under different operating conditions is:

[0032]

[0033] In the formula, q t represents the radiation heat flux density, ω i represents the weight, τχ r m represents the heat source parameter, S Δ Indicates the distribution of different distances between the heat source and the radiated object; θ represents the angle between the normal of the reflecting surface of the quartz lamp array and the heat source-structure, θ Δ Indicates the angle distribution between the normal direction of the reflecting surface of the quartz lamp array and the heat source-structure; ΔH cIt represents the heat of combustion per unit mass of fuel.

[0034] Furthermore, in step 4, the heat capacity of the heat flow test device is corrected by adjusting q t The weight of the calculation process ω i , heat source parameter τχ r m is adjusted so that the error between the measured and calculated values ​​of the gas heat flow under different engine operating conditions is controlled within 1%.

[0035] Furthermore, in step 5, the calculation formula for the heat flux radiated outward from the sensitive layer is:

[0036] Q4=εσAT 4

[0037] Where ε is the emissivity of the sensitive layer, which refers to the ratio of the energy radiated by the object itself to the energy radiated by an absolute black body at the same temperature. The emissivity is only related to the properties of the object's surface; σ is the Boltzmann constant, A is the area of ​​the sensitive layer, and T is the temperature of the sensitive layer under different working conditions.

[0038] Furthermore, in step 6, the calculation formula of the heat flux gradient is:

[0039]

[0040] Among them, c p is the equivalent heat capacity of the sensitive layer, m is the mass of the sensitive layer, T is the temperature of the sensitive layer, and t is the time; Q1 is the heat flux emitted by the quartz lamp array and absorbed by the sensitive layer, that is, the heat flux q of the quartz lamp array, q = εσT 4 / a, where ε is the surface emissivity of the sensitive layer, which refers to the ratio of the energy radiated by the object itself to the energy radiated by an absolute black body at the same temperature. The emissivity is only related to the properties of the object surface; a is the absorptivity of the sensitive layer; σ is the Boltzmann constant, which is 5.67×10 -8 W / (m 2 ·K 4 ), T is the surface temperature of the sensitive layer under different working conditions.

[0041] Compared with the prior art, the present invention has the following beneficial effects:

[0042] 1. The heat flow testing device of the present invention includes a fixed base, a displacement adjustment base and an angle adjustment base respectively disposed below the fixed base, and a thermal insulation layer, a temperature sensor, and a sensitive layer disposed sequentially above the fixed base. The displacement adjustment base is used to enable movement of the fixed base during measurement, and the angle adjustment base is used to adjust the installation angle of the fixed base during measurement, thereby enabling real-time measurement of heat flow at different engine locations. The sensitive layer serves as the plume receiving surface and is coated above the temperature sensor, which is in turn located above the thermal insulation layer. This design allows the temperature changes of the sensitive layer to be captured in real time by the temperature sensor. The thermal insulation layer also prevents heat transfer from the sensitive layer to the surrounding area, thereby improving measurement accuracy and reducing measurement errors.

[0043] 2. The test method of the present invention obtains heat flux and thermal effect data of different scales under the ignition state of the attitude control engine by inverting the real plume thermal effect data of the attitude control engine on orbit, combining the radiation heat flux and external radiation heat flux of the quartz lamp array, and realizing the measurement of the heat flux gradient of the thermal environment at the moment of engine ignition, ensuring the continuity and reliability of the heat flux loading during the test, and providing a guarantee for verifying the working performance of the main power engine of the aircraft in the working state and under the simultaneous action of external heat flux. BRIEF DESCRIPTION OF THE DRAWINGS

[0044] Figure 1 This is a schematic structural diagram of an embodiment of a heat flow test device for simulating heat flow in a flight environment of an attitude control engine according to the present invention;

[0045] Figure 2 A reference diagram of the test process of an embodiment of a test method for simulating heat flow in a flight environment of an attitude control engine according to the present invention;

[0046] Figure 3 This is a position distribution diagram of the heat flow gradient test in an embodiment of the test method for simulating heat flow in the attitude control engine flight environment of the present invention.

[0047] The reference numerals are as follows:

[0048] 1-Fixed seat, 2-Displacement adjustment seat, 3-Angle adjustment, 4-Thermal insulation layer, 5-Temperature sensor, 6-Sensitive layer. DETAILED DESCRIPTION

[0049] The present invention is further described in detail below with reference to the accompanying drawings and specific embodiments.

[0050] The design principle of the present invention is: for the mechanical and thermal coupling instantaneous thermal environment formed by engine ignition, by analyzing the radiation heat exchange relationship between the attitude control engine structure and the infrared heating lamp array and temperature, while keeping the physical parameters and position relationship unchanged, the difference in the consistency of the equivalent absorbed heat flux density on the engine surface is analyzed to obtain the relationship between the heat flux density and the current, and the factors affecting its unevenness are analyzed. The optimized power-on method of the infrared lamp array heating under transient working conditions is determined, and a heat flow calculation model under the engine ignition jet state is established to reduce the transient external heat flow simulation error of the infrared lamp array and achieve uniformity and accurate measurement of the instantaneous heat flow loading in the mechanical and thermal coupling environment.

[0051] The present invention provides a heat flow test device for simulating heat flow in the flight environment of an attitude control engine. The heat flow test device is installed on the outer side of the nozzle which is significantly affected by the attitude control engine. Figure 1 As shown, it includes a fixed base 1, a displacement adjustment base 2 and an angle adjustment base 3 respectively arranged below the fixed base 1, and a heat insulation layer 4, a temperature sensor 5 and a sensitive layer 6 arranged in sequence above the fixed base 1. Among them, the displacement adjustment base 2 is used to adjust the position of the test device during the measurement process; the angle adjustment base 3 is used to adjust the measurement angle.

[0052] Thermal insulation layer 4 covers the top of mounting base 1. As the temperature of sensitive layer 6 rises, it transfers heat to the surrounding area, affecting the measurement results. To accurately measure the plume value, it is necessary to convert as much plume heat as possible into a temperature rise in sensitive layer 6, minimizing heat transfer from the sensitive layer 6 to the surrounding area. Therefore, a multi-layered thermal insulation layer 4 is used for thermal insulation.

[0053] The temperature sensor 5 is located above the thermal insulation layer 4; the sensitive layer 6 is wrapped around the temperature sensor 5 and faces the attitude control engine. Stainless steel wire is used to bundle the two sides of the temperature sensor 5 and is fixed with silicone rubber to assist in the fixation, so that the temperature sensor 5 can directly sense the temperature changes of the sensitive layer 6.

[0054] Sensitive layer 6 is the plume-receiving surface. Its function is to absorb the plume from the high-thrust engine and convert it into a change in its own temperature, which is then detected by temperature sensor 5. To increase the sensitivity of sensitive layer 6, its thickness should be as thin as possible, generally controlled at 50 to 100 μm. Furthermore, sensitive layer 6 needs to have a certain mechanical strength to secure temperature sensor 5, so stainless steel foil is selected as the sensitive layer. To ensure the stability of the thermal flow test device, a high-temperature resistant and anti-oxidation coating is applied above sensitive layer 6. During testing, the side of sensitive layer 6 with the high-temperature resistant and anti-oxidation coating is facing the attitude control engine.

[0055] In order to improve the temperature sensitivity of the heat flow test device, the heat capacity of the heat flow test device should be as small as possible. During installation, first cover the fixing seat 1 with a heat insulation layer 4, and then stick the temperature sensor 5 for temperature measurement on the outer surface of the heat insulation layer 4; then cover the outside of the temperature measuring point with a layer of stainless steel foil with a high-temperature anti-oxidation coating, and tie the two sides of the temperature measuring point with stainless steel wire to ensure that the stainless steel foil is in contact with the temperature measuring point. The heat flow test device can monitor the plume thermal effect value of the attitude control power engine when it is working in orbit. It can adapt to a wide temperature range of 1000℃ and 400KW / m 2 Effective monitoring of the thermal effects of short-duration ultra-high plumes.

[0056] In addition, the present invention also provides a test method for simulating heat flow in the flight environment of an attitude control engine, such as Figure 2 As shown, the specific steps include:

[0057] Step 1, such as Figure 3 As shown in FIG, a quartz lamp array is used to simulate the gas jet process of an attitude control engine in a flight environment, and a heat flux test device is used to measure the radiation heat flux Q2 of the quartz lamp array at different heights, different spacings and different voltages.

[0058] Quartz lamp radiant heating is a widely used heating method in aircraft structural thermal testing. Its greatest advantage is its low thermal inertia, which makes it easy to control and assemble into heaters of various shapes to suit the test specimen's appearance. Using the Monte Carlo method, which assumes that light beams carry no energy, after tracking a large number of light beam transfers, a probabilistic simulation is used to calculate the radiation transfer factor. The radiation transfer factor is defined as the fraction of radiation energy ultimately absorbed by a unit in a radiation transmission system after a single projection and one or more reflections from other units within the system. In this embodiment, the calculation method for the radiation heat flux Q2 of the quartz lamp array at different heights, different spacings, and different voltages is as follows:

[0059]

[0060] Where: q s (x, y, z) is the radiation heat flux Q2, which is the radiation heat flux at the coordinate (x, y, z) on the quartz lamp plate. Q0 is the emission power of the quartz lamp, and H is the distance between the center of the quartz lamp and the plate. L is the heating length of the quartz lamp filament; α, β, and γ are all custom values, and is the angle between the heat flow measurement device and the engine axis.

[0061] Step 2: Based on the radiation heat flux Q2 in step 1, the uniformity of the heat flux loading is obtained, and then the sensitivity of the heat flux loading is obtained.

[0062] In this embodiment, when using a quartz lamp array to simulate external heat flux, non-uniformity is an important parameter used to evaluate the degree of uniformity of heat flux distribution on the illuminated surface (or the illuminated area of ​​interest during the test). The heat flux loading non-uniformity is:

[0063]

[0064] Where, δ is the unevenness of the quartz lamp array on the illuminated surface; q max is the maximum heat flux density on the illuminated surface; q min is the minimum heat flux density on the illuminated surface.

[0065] pass It can be seen that when the maximum and minimum heat flux densities are known, the heat flux nonuniformity can be determined. Within the surface of the specimen in the area illuminated by the infrared lamp array, the heat flux density is measured at a sufficient number of points, and the finite number of measured heat flux values ​​is ultimately used to characterize the heat flux distribution on the illuminated surface. The relationship between the transient arrival heat flux density of the environment, the equivalent absorbed heat flux density of the engine surface, and the current applied to the lamp array is derived. The maximum and minimum heat flux densities are then taken to determine the nonuniformity of the lamp array at that time. To simulate conditions under different heat flux densities, the current of the quartz lamp needs to be varied. Different currents may result in different nonuniformity δ.

[0066] At the same time, combined with different influencing factors (distance, angle, working environment, vacuum degree, lamp array power), Obtain the induction coefficients of different influencing factors, thereby obtaining the sensitivity of heat flow loading;

[0067]

[0068] in: The sensitivity of the heat flow loading to a certain influencing factor.

[0069] Step 3: Use a heat flow test device to perform heat flow tests on the engine in different working conditions and at different parts in the axial and radial directions to obtain the heat flow q of the engine in different working conditions. t , and the rate of change of the axial and radial heat flux gradients of the engine under different operating conditions.

[0070] There are currently two main research methods for modeling the thermal radiation distribution around a jet heat flow: one is a numerical simulation method based on computational fluid dynamics theory, and the other is a semi-empirical model prediction method based on basic theoretical assumptions. The single-point source model was first proposed by Mudan et al. to predict pool fire thermal radiation. Based on this model, assuming that all radiant heat is caused by this heat source, the heat flux distribution correction model is:

[0071]

[0072] Where q t Indicates the radiation heat flux density in W / m 2 , τχ r m represents the heat source parameter, which is related to the engine tail flame properties, S Δ It represents the distribution of different distances between the heat source and the radiated object, with the unit being m. θ represents the distribution of different angles between the normal of the reflecting surface and the heat source-structure. ΔH c The heat of combustion per unit mass of fuel is expressed in J / kg. Based on the single-point source model, a thermal radiation prediction model using weighted superposition is proposed. This model discretizes the continuous flame into multiple heat sources, where the radiant heat at any location is the linear sum of these multiple heat sources.

[0073] In this embodiment, different operating conditions generally refer to different propellant flow rates, different vacuum levels, different flight altitudes, different incident angles, different quartz lamp array powers or voltages, different reflection coefficients, etc. The calculation formula for the combustion gas heat flow of the engine under different operating conditions is:

[0074]

[0075] In the formula, q t represents the radiation heat flux density, ω i represents the weight, τχ r m represents the heat source parameter, S Δ Indicates the distribution of different distances between the heat source and the radiated object, θ represents the angle between the normal of the reflecting surface of the quartz lamp array and the heat source-structure, θ Δ Indicates the angle distribution between the normal direction of the reflecting surface of the quartz lamp array and the heat source-structure; ΔH c It represents the heat of combustion per unit mass of fuel.

[0076] Assuming that the jet direction is along the straight line y=ax+b, then any point source at the above position satisfies (x,y)=(x,ax+b), k represents the positive correlation coefficient with each heat source. Given the (x,y,z) array and the number of heat sources, the least squares method is used to fit the parameters in the equation to determine the heat source and the jet direction, and then the distribution of the radiation heat flux gradient can be determined.

[0077] Step 4: Compare the engine gas heat flux under different operating conditions obtained in step 3 with the gas heat flux value during actual flight. If the error between the two is within 1% to 5%, proceed to step 5; otherwise, correct the heat capacity value of the heat flux testing device and return to step 1.

[0078] In this embodiment, the correction method for the heat capacity of the heat flow test device is: t The weight of the calculation process (ω i ), heat source parameter τχ rm is adjusted so that the error between the measured and calculated values ​​of the gas heat flow under different engine operating conditions is controlled within 1%.

[0079] Step 5: Test the heat flux output value of the engine under different working conditions using a heat flux testing device to obtain the reflectivity of the heat flux testing device, thereby obtaining the outward radiation heat flux Q4 of the sensitive layer under different working conditions.

[0080] In this embodiment, in step 5, the calculation formula for the heat flux radiated outward from the sensitive layer is:

[0081] Q4=εσAT 4

[0082] Where ε is the emissivity of the sensitive layer, σ is the Boltzmann constant, A is the area of ​​the sensitive layer, and T is the temperature of the sensitive layer under different working conditions.

[0083] Step 6: By inverting the actual on-orbit heat flux data of the attitude control engine, combined with the quartz lamp array's radiant heat flux Q2 obtained in Step 1, the outward radiant heat flux Q4 of the sensitive layer obtained in Step 5, the heat flux Q1 absorbed by the sensitive layer from the infrared heat flux emitted by the quartz lamp array, and the conductive heat flux Q3 of the sensitive layer, the heat flux gradient at the instant of attitude control engine ignition is calculated. Due to the presence of the thermal insulation layer, the conductive heat flux Q3 of the sensitive layer can generally be ignored.

[0084] In this embodiment, in step 6, due to the low air pressure of the simulated external environment at high altitude, the high-altitude engine jet flow field expands and reverses, thereby forming convective heating on the bottom solid surface. The heat flux gradient is calculated as follows:

[0085]

[0086] Among them, c p is the equivalent heat capacity of the sensitive layer, m is the mass of the sensitive layer, T is the temperature of the sensitive layer under different working conditions, and t is time; Q1 is the part of the infrared heat flux emitted by the quartz lamp array that is absorbed by the sensitive layer, that is, the arriving heat flux q of the quartz lamp array, q=εσT 4 / a, where ε is the surface emissivity of the sensitive layer, a is the absorptivity of the sensitive layer; σ is the Boltzmann constant, which is 5.67×10 -8 W / (m 2 ·K 4 ), T is the surface temperature of the sensitive layer under different working conditions, the unit is K.

[0087] Finally, in order to analyze the effects of different engine operating conditions, propellant flow rates, vacuum levels, flight altitudes, incident angles, quartz lamp array power or voltage, and reflection coefficients on heat flux measurement, and to conduct precise measurement and evaluation of heat flux gradients under multi-factor conditions in the thermal environment at the moment of engine ignition, the present invention proposes the following definitions:

[0088] 1) Heat flow sensitivity: This refers to the degree to which factors such as different engine operating conditions, different propellant flow rates, different vacuum levels, different test positions, different incident angles, different quartz lamp array power or voltage, and emissivity directly affect the heat flow gradient value measured by the heat flow measurement device.

[0089] 2) Gradient change rate: It is divided into axial gradient change rate and radial gradient change rate, which indicates the different heat flux distribution and changes obtained by the heat flux test device due to different axial positions and radial positions;

[0090] 3) Coupling degree: It indicates the degree of influence of the mutual influence or connection between different engine operating conditions, different propellant flow rates, different vacuum levels, different test positions, different incident angles, different quartz lamp array power or voltage, reflectivity and other factors on the heat flux gradient value;

[0091] 4) Matching degree: Indicates the adaptability and effectiveness of the heat flow measurement device to different vacuum simulations, different types of engine ignition operations, different flight speeds, different propellant compositions, different power lamp array heating and other environmental changes.

[0092] The present invention is based on the accurate measurement process of heat flow under the engine hot gas jet state, realizes the thermal analysis model of the engine ignition state under different working conditions, and obtains the heat flow thermal effect data of the attitude control engine under the ignition state by inverting the real plume thermal effect data of the attitude control engine on orbit, combining the radiation heat flux and the external radiation heat flux of the quartz lamp array, and realizes the measurement of the thermal flow gradient of the thermal environment at the moment of engine ignition. This method improves the measurement accuracy through multi-gradient measurement, thereby providing theoretical support for the actual flight environment of the attitude control engine.

[0093] Although the embodiments of the present invention have been shown and described above, it will be apparent to those skilled in the art that any changes or modifications to the above embodiments shall fall within the scope of protection of the present invention as long as they are within the spirit of the present invention.

Claims

1. A thermal flow test device for simulating thermal flow in the flight environment of an attitude control engine, which is used to be installed outside a nozzle affected by the attitude control engine, and is characterized by: It comprises a fixed seat (1), a displacement adjustment seat (2) and an angle adjustment seat (3) respectively arranged below the fixed seat (1), and a heat insulation layer (4), a temperature sensor (5) and a sensitive layer (6) arranged above the fixed seat (1); The displacement adjustment seat (2) is used to adjust the measurement position, and the angle adjustment seat (3) is used to adjust the measurement angle; The heat insulation layer (4) covers the top of the fixing seat (1), and the heat insulation layer (4) includes a plurality of heat insulation layers covered in sequence, and is used to reduce the heat conduction of the sensitive layer (6) to the surroundings after absorbing the heat flow; The temperature sensor (5) is located above the heat insulation layer (4); the sensitive layer (6) is coated above the temperature sensor (5) and faces the attitude control engine, and is used to absorb the heat flow of the engine and convert it into its own temperature. The temperature sensor (5) then collects this temperature information and outputs it to an external device.

2. The thermal flow testing device for simulating thermal flow in the flight environment of an attitude control engine according to claim 1, characterized in that: The thickness of the sensitive layer (6) is 50-100 μm, and a high-temperature resistant and anti-oxidation coating is provided above the sensitive layer (6), and the high-temperature resistant and anti-oxidation coating faces the attitude control engine.

3. The thermal flow testing device for simulating thermal flow in the flight environment of an attitude control engine according to claim 2, characterized in that: The material of the sensitive layer (6) is stainless steel foil; The heat insulation layer (4) is formed by laminating double-sided aluminum-plated polyimide films.

4. A test method for simulating heat flow in an attitude control engine flight environment, characterized in that: The thermal flow testing device for simulating thermal flow in a flight environment of an attitude control engine according to any one of claims 1 to 3 comprises the following steps: Step 1: Use a quartz lamp array to simulate the gas jet process of an attitude control engine in a flight environment, and use a heat flux test device to measure the radiation heat flux Q2 of the quartz lamp array at different heights, different spacings, and different voltages; Step 2: Based on the radiation heat flux Q2 measured in step 1, the uniformity distribution of the heat flux loading is obtained, and then the sensitivity of the heat flux loading is obtained; Step 3: Use a heat flow test device to perform heat flow tests on the engine in different working conditions and at different parts in the axial and radial directions to obtain the heat flow q of the engine in different working conditions. t , and the rate of change of axial and radial heat flux gradients of the engine under different operating conditions; Step 4: Compare the engine gas heat flux under different operating conditions obtained in Step 3 with the gas heat flux value during flight. If the error between the two is within 1% to 5%, proceed to Step 5; otherwise, correct the heat capacity value of the heat flux testing device and return to Step 1. Step 5: Test the heat flux output value of the engine under different working conditions using a heat flux testing device to obtain the reflectivity of the heat flux testing device, thereby obtaining the outward radiation heat flux Q4 of the sensitive layer under different working conditions; In step 6, the heat flux gradient in the thermal environment at the moment of attitude control engine ignition is calculated by inverting the actual heat flux data of the attitude control engine on orbit, combining the radiation heat flux Q2 of the quartz lamp array obtained in step 1 and the outward radiation heat flux Q4 of the sensitive layer obtained in step 5, as well as the heat flux Q1 of the infrared heat flux emitted by the quartz lamp array and absorbed by the sensitive layer and the conduction heat flux Q3 of the sensitive layer.

5. The test method for simulating heat flow in the attitude control engine flight environment according to claim 4, characterized in that: In step 1, the calculation method of the radiation heat flux Q2 of the quartz lamp array at different heights, different spacings and different voltages is: Where: q s (x, y, z) is the radiation heat flux Q2, which is the radiation heat flux at the coordinate (x, y, z) on the plate, Q0 is the quartz lamp emission power, H is the distance between the center of the quartz lamp and the plate, L is the heating length of the quartz lamp filament; α, β, and γ are all custom values, and is the angle between the heat flow measurement device and the engine axis.

6. The method for testing thermal flow in a simulated flight environment of an attitude control engine according to claim 5, characterized in that: Specifically, step 2 is to obtain the uniformity distribution of the heat flux loading based on the radiation heat flux Q2 measured in step 1, and then obtain the non-uniformity of the heat flux loading. The non-uniformity calculation formula is as follows: Where, δ is the unevenness of the quartz lamp array on the illuminated surface; q max is the maximum heat flux density on the illuminated surface; q min is the minimum heat flux density on the illuminated surface; At the same time, the induction coefficients of different influencing factors are obtained by combining different influencing factors, thereby obtaining the sensitivity of heat flux loading; the different influencing factors include the distribution distance, angle, working environment and vacuum degree of the quartz lamp array; the sensitivity of the heat flux loading Expressed as: in: That is, it represents the sensitivity of heat flux loading under different influencing factors.

7. The method for testing thermal flow of an attitude control engine in a simulated flight environment according to claim 6, characterized in that: In step 3, the calculation formula for the gas heat flow of the engine under different operating conditions is: In the formula, q t represents the radiation heat flux density, ω i represents the weight, τχ r m represents the heat source parameter, S Δ Indicates the distribution of different distances between the heat source and the radiated object, θ represents the angle between the normal of the reflecting surface of the quartz lamp array and the heat source-structure, θ Δ Indicates the angle distribution between the normal direction of the reflecting surface of the quartz lamp array and the heat source-structure; ΔH c It represents the heat of combustion per unit mass of fuel.

8. The method for testing thermal flow of attitude control engine flight environment simulation according to claim 7, characterized in that: In step 4, the heat capacity of the heat flow test device is corrected by adjusting q t The weight of the calculation process ω i , heat source parameter τχ r m is adjusted so that the error between the measured and calculated values ​​of the gas heat flow under different engine operating conditions is controlled within 1%.

9. The method for testing thermal flow of an attitude control engine in a simulated flight environment according to claim 8, characterized in that: In step 5, the calculation formula for the heat flux radiated outward from the sensitive layer is: Q4=εσAT 4 Wherein, ε is the surface emissivity of the sensitive layer; σ is the Boltzmann constant; A is the area of ​​the sensitive layer; and T is the temperature of the sensitive layer under different working conditions.

10. The method for testing thermal flow of attitude control engine flight environment simulation according to claim 9, characterized in that: In step 6, the heat flux gradient is calculated as follows: Among them, c p is the equivalent heat capacity of the sensitive layer, m is the mass of the sensitive layer, T is the temperature of the sensitive layer under different working conditions, and t is the time; Q1 is the heat flux of the infrared heat flux emitted by the quartz lamp array and absorbed by the sensitive layer, that is, the heat flux q of the quartz lamp array, q = εσT 4 / a, where ε is the surface emissivity of the sensitive layer, a is the absorptivity of the sensitive layer; σ is the Boltzmann constant, which is 5.67×10 -8 W / (m 2 ·K 4 ).

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