Liquid rocket engine test stand exhaust energy efficient utilization system

Abstract

The application belongs to the technical field of liquid rocket engine test, and relates to a liquid rocket engine test bench tail gas energy efficient utilization system, which comprises an air source, an adjustable throat area nozzle assembly, a turbine power generation device and a controller. The air source is connected with one end of the adjustable throat area nozzle assembly through a first pipeline, and the other end of the adjustable throat area nozzle assembly is connected with the turbine power generation device through a second pipeline. The controller is connected with the first pipeline, the second pipeline and the adjustable throat area nozzle assembly to control the throat area change of the adjustable throat area nozzle assembly according to monitored pipeline parameters, and then adjust the nozzle outlet velocity, so that the turbine work can be at a higher working efficiency point. The application can improve the power generation efficiency of the power generation device.

Description

TECHNICAL FIELD

[0001] The present application relates to the technical field of liquid rocket engine test, in particular to a liquid rocket engine test bench tail gas energy efficient utilization system. BACKGROUND

[0002] Low-temperature non-toxic propellant has strong competitiveness in the international rocket commercial market, and its development mainly depends on its two characteristics: high performance and low cost. With the increasing demand of space and national defense, especially the long-term on-orbit deep space exploration flight and the reusable earth-to-space flight, liquid rocket engines represented by liquid oxygen / liquid hydrogen and liquid oxygen / methane have gradually attracted attention.

[0003] The development of low-temperature liquid rocket engines requires a large number of ground tests. Hydrogen and methane, as two kinds of flammable and explosive propellants, need to be blown off before and after the test to ensure safety. Unlike normal temperature propellants, low-temperature propellants have very low boiling points. At standard atmospheric pressure, the boiling point of liquid hydrogen is only 20.38 K, and the boiling point of liquid oxygen is 90.15 K. They are in gaseous state at room temperature, which is easy to vaporize and appear gas-liquid two-phase flow during storage and transportation, causing adverse or even destructive effects on the engine and pressurized delivery system. Complete ground thermal test includes propellant pipeline cleaning, gas blowing, and pipeline pre-cooling, which requires high-pressure gas as a power source. After completing the above processes, most of the mechanical energy of the gas is not fully released, but is directly discharged into the atmosphere. Therefore, how to optimize the process and recycle this part of energy is of great significance.

[0004] According to relevant data, from the energy quality point of view, only 20% to 45% of the total heat energy generated by the combustion of fuel in the cylinder of the automobile engine is converted into effective mechanical work, and the remaining heat energy is lost in the form of exhaust gas, cooling medium and surface heat dissipation of parts, among which the energy carried away by the exhaust gas accounts for 30% to 40% of the total energy. For the gas blowing and pre-cooling measures in the ground test process of the rocket engine, only a small part of the mechanical energy of the high-pressure energy gas is lost due to the flow resistance of the pipeline.

[0005] In the prior art, the main ways to realize the energy recovery of engine exhaust gas include three forms of turbo compound technology, Rankine cycle technology, and thermoelectric exchange technology.

[0006] Among these technologies, turbine-integrated turbine technology, particularly turbine-powered turbine, utilizes a combined cycle with a turbine to recover and utilize the mechanical energy of high-pressure gas. The high-pressure exhaust gas discharged through the pipeline drives a power turbine installed in the exhaust pipe, causing it to rotate at high speed. This converts the energy contained in the exhaust gas into the mechanical energy of the power turbine. The high-speed rotating power turbine then drives a generator to charge batteries, demonstrating promising engineering application prospects. The implementation and system structure of turbine-integrated turbine technology are relatively simple, requiring minimal modifications to the engine structure and offering flexible control, thus attracting widespread attention in the industry. Turbine-integrated turbine technology is of great significance for engine exhaust gas energy recovery and has considerable potential. Furthermore, the combination of turbines and engines is not limited to a single form. The main combinations of turbines and engines include electrically assisted composite turbines, series composite turbines, parallel composite turbines, and pure electric composite turbines.

[0007] However, the turbine structure in current turbine power generation systems is fixed, which means that the turbine does not operate at its design point most of the time, resulting in a low efficiency level and consequently, low power generation efficiency for the entire turbine power generation unit. Adapting the turbine to the upstream flow conditions is an effective way to improve the power generation efficiency of turbine power generation systems. Summary of the Invention

[0008] Therefore, it is necessary to provide a system for efficient utilization of exhaust gas energy from a liquid rocket engine test rig to address the aforementioned technical problems and improve the power generation efficiency of the power generation device.

[0009] The liquid rocket engine test rig exhaust gas energy high-efficiency utilization system includes: gas source, adjustable throat area nozzle assembly, turbine power generation device and controller;

[0010] The gas source is connected to one end of the adjustable throat area nozzle assembly via a first pipeline, and the other end of the adjustable throat area nozzle assembly is connected to the turbine power generation device via a second pipeline.

[0011] The controller is connected to the first pipeline, the second pipeline, and the adjustable throat area nozzle assembly, respectively, to control the throat area change of the adjustable throat area nozzle assembly according to the monitored pipeline parameters.

[0012] In one embodiment, the adjustable throat area nozzle assembly includes: a motor, a needle cone, and a nozzle;

[0013] The motor is detachably connected to one end of the needle cone;

[0014] The central axis of the needle cone is collinear with the center of the throat of the nozzle, and the other end of the needle cone has a gap with the throat of the nozzle;

[0015] The controller is connected with the motor and outputs control signals with the pipeline parameters as input, so as to control the motor to move the control needle cone in the direction towards or away from the throat of the nozzle, so as to change the throat area of the adjustable throat area nozzle assembly.

[0016] In one embodiment, the nozzle comprises a converging section, a throat and a diverging section connected in sequence.

[0017] In one embodiment, the converging section adopts a double-arc design.

[0018] In one embodiment, the diverging section adopts a special nozzle design.

[0019] In one embodiment, the first pipeline is provided with a temperature sensor and a pressure sensor.

[0020] The temperature sensor and the pressure sensor are connected with the controller, so that the controller controls the throat of the nozzle according to the temperature monitored by the temperature sensor and the pressure monitored by the pressure sensor.

[0021] In one embodiment, the turbine power generation device comprises a turbine moving blade, a transmission mechanism and a generator.

[0022] One end of the turbine moving blade is connected with the nozzle, and the other end is connected with the generator through the transmission mechanism.

[0023] In one embodiment, the turbine moving blade adopts a single-stage impulse turbine.

[0024] In one embodiment, the isentropic expansion work of the working medium of the turbine moving blade in the single-stage turbine satisfies the equation:

[0025]

[0026] In the formula, L ad is the isentropic expansion work, Ω is the turbine reaction rate, R is the universal gas constant, k is the specific heat ratio of the turbine working medium, is the total temperature at the turbine inlet, π is the isentropic expansion work of the turbine gas in the nozzle.

[0027] In one embodiment, the tangential velocity of the airflow at the average diameter of the turbine moving blade satisfies:

[0028]

[0029] In the formula, is the tangential velocity of the airflow at the average diameter, is the actual velocity at the nozzle outlet, is the nozzle exit angle.

[0030] The liquid rocket engine test bench tail gas energy efficient utilization system adjusts the nozzle throat area, and then adjusts the nozzle expansion ratio, and then adjusts the nozzle outlet flow velocity, so as to realize the adjustment of the nozzle outlet velocity under different upstream conditions, so that the turbine moving blade always works at a high efficiency level; through the sensing of the pressure and temperature sensors and the mutual cooperation of the PID controller, the movement of the adjustable nozzle needle cone is realized, and then the intelligent adjustment of the device is realized; through the efficient design of the turbine moving blade and the mutual cooperation of the turbine moving blade and the adjustable nozzle, the efficient operation of the whole turbine device is realized. BRIEF DESCRIPTION OF DRAWINGS

[0031] Figure 1 is a schematic diagram of the liquid rocket engine test bench tail gas energy efficient utilization system in an embodiment;

[0032] Figure 2 is a sectional view of the nozzle in the liquid rocket engine test bench tail gas energy efficient utilization system in an embodiment;

[0033] Figure 3 is a schematic diagram of the nozzle convergent section in the liquid rocket engine test bench tail gas energy efficient utilization system in an embodiment;

[0034] Figure 4 is a schematic diagram of the nozzle divergent section in the liquid rocket engine test bench tail gas energy efficient utilization system in an embodiment;

[0035] Figure 5 is a control principle diagram of the liquid rocket engine test bench tail gas energy efficient utilization system in an embodiment. DETAILED DESCRIPTION

[0036] In order to make the purpose, technical scheme and advantages of the present application clearer, the present application will be further described in detail below in combination with the drawings and embodiments. It should be understood that the specific embodiments described herein are only used to explain the present application and not to limit the present application. Based on the embodiments in the present application, all other embodiments obtained by those of ordinary skill in the art without creative labor fall within the scope of the present application.

[0037] It should be noted that all directional indications (such as up, down, left, right, front, back, etc.) in the embodiments of the present application are only used to explain the relative positional relationship, movement condition, etc. between components in a certain posture (as shown in the drawings), and if the certain posture changes, the directional indications also change accordingly.

[0038] In addition, the descriptions such as "first", "second" and the like in the present application are only for the purpose of description, and cannot be understood as indicating or implying the relative importance of the technical features indicated or implying the number of the technical features indicated. Therefore, the features defined as "first" and "second" can be explicitly or implicitly included at least one of the features. In the description of the present application, the meaning of "multiple groups" is at least two groups, such as two groups, three groups, etc., unless otherwise specifically limited.

[0039] In the present application, unless otherwise specifically defined and limited, the terms "connection", "fixing" and the like should be understood in a broad sense, for example, "fixing" can be fixed connection, or detachable connection, or integral; can be mechanical connection, or electrical connection, or physical connection, or wireless communication connection; can be directly connected, or indirectly connected through intermediate medium, can be internal communication of two elements or interaction relationship between two elements, unless otherwise specifically limited. For those skilled in the art, the specific meaning of the above terms in the present application can be understood according to the specific circumstances.

[0040] In addition, the technical solutions of various embodiments of the present application can be combined with each other, but it must be based on the fact that the ordinary skilled in the art can realize it, when the combination of technical solutions appears contradictory or unachievable, it should be considered that the combination of technical solutions does not exist, nor within the scope of protection claimed in the present application.

[0041] The present application provides a liquid rocket engine test bench tail gas energy efficient utilization system, as shown in Figure 1 In one embodiment, it comprises: a gas source, a first pipeline, an adjustable throat area nozzle assembly, a second pipeline, a turbine power generation device and a controller.

[0042] The gas source is connected to one end of the adjustable throat area nozzle assembly through the first pipeline, and the other end of the adjustable throat area nozzle is connected to the turbine power generation device through the second pipeline.

[0043] The controller is connected to the first pipeline, the second pipeline and the adjustable throat area nozzle assembly respectively, so as to control the change of the throat area of the adjustable throat area nozzle assembly according to the monitored pipeline parameters.

[0044] The first pipeline is provided with a temperature sensor and a pressure sensor; the temperature sensor and the pressure sensor are connected to the controller, so that the controller controls the throat of the adjustable throat area nozzle assembly according to the temperature monitored by the temperature sensor and the pressure monitored by the pressure sensor.

[0045] As shown in Figure 2As shown, in one embodiment, the adjustable throat area nozzle assembly comprises: a motor 1, a needle cone 2 and a nozzle 3.

[0046] The motor is detachably connected to one end of the needle cone, such as a threaded connection 4, and the motor can be a stepper motor.

[0047] The central axis of the needle cone is collinear with the throat center of the nozzle, and the other end of the needle cone has a gap with the throat of the nozzle, and 5 is the diameter of the needle cone.

[0048] The nozzle is a fixed profile Laval nozzle.

[0049] The controller is connected to the motor and outputs a control signal with the pipeline parameters as input, and moves the needle cone in the direction towards or away from the nozzle throat by controlling the motor, so as to change the throat area of the adjustable throat area nozzle assembly.

[0050] As shown Figure 3 And Figure 4 Because the exhaust gas generated during the test process is often high in pressure, in order to simplify the structure and reduce the processing cost, a circular convergent-divergent nozzle, i.e. a Laval nozzle commonly used in liquid rocket engines, is adopted. The nozzle comprises a convergent section, a throat and a divergent section connected in sequence. The convergent section adopts a double-arc design, the length of the cylindrical section is L c1 , the length of the convergent section is L c2 , the nozzle throat radius is R t , the circular arc radius of the part connecting the nozzle inlet and the combustion chamber is R2, i.e. the throat circular arc, and the circular arc with a radius of R1 is tangent to the throat circular arc. The divergent section adopts a double-circular-arc special nozzle design, the nozzle divergent section length is L c3 , the nozzle throat radius is R t , and the nozzle outlet radius is R e . The nozzle throat profile is a nozzle with a radius of R3, and the divergent section is a circular arc with a radius of R4, and the two circular arcs are tangent at point e , and β e is the nozzle outlet expansion angle.

[0051] In one embodiment, the turbine device comprises: turbine moving blades, a transmission mechanism and a generator; for small power turbines, the structure is simple, the process is good, and the work is reliable, so the turbine moving blades adopt single-stage impact turbines, one end of the turbine moving blades is connected to the nozzle, and the other end is connected to the generator through the transmission mechanism. The reaction degree of the turbine device is 0, i.e. the expansion of the gas in the turbine device only occurs in the nozzle assembly. Since the design rotating speed of the turbine is generally high, in order to match the allowable rotating speed range of the generator with the rotating speed of the turbine, a transmission device is installed between the turbine and the generator.

[0052] As shown Figure 5As shown, the working process of the present application is as follows: after the liquid rocket engine test is completed, the test exhaust gas is blown out as a gas source from the upstream, and at the same time the high-pressure electromagnetic valve is opened. The gas flows through the adjustable throat area nozzle assembly and the turbine moving blade and is discharged into the environment, and then the moving blade rotates, the moving blade drives the transmission mechanism to rotate, and then drives the generator to rotate to generate electric energy. When the upstream inlet parameter changes, the temperature and pressure sensors along the pipeline measure the inlet parameter, and the changing parameter is input into the PID controller. The PID controller outputs a control signal to control the stepping motor to move the needle cone, and adjusts the nozzle throat area to make the gas flow expand in the nozzle, so that the outlet pressure is equal to the ambient pressure, and then the gas is fully expanded, the nozzle outlet gas velocity is improved, and the power generation efficiency of the turbine power generation device is also improved.

[0053] The liquid rocket engine test bench tail gas energy efficient utilization system adjusts the adjustable nozzle throat area, and then adjusts the nozzle expansion ratio, and then adjusts the nozzle outlet gas flow velocity, to realize the adjustment of the nozzle outlet velocity under different upstream conditions, so that the turbine moving blade always works at a high efficiency level. Through the sensing of the pressure and temperature sensors and the cooperation of the PID controller, the movement of the adjustable nozzle needle cone is realized, and then the intelligent adjustment of the device is realized. Through the efficient design of the turbine moving blade and the cooperation of the turbine moving blade and the adjustable nozzle, the efficient operation of the whole turbine power generation device is realized.

[0054] Compared with the traditional fixed structure turbine power generation device, the present application has higher efficiency, simple structure, easy realization and lower cost. In addition, in addition to the test device tail gas energy recovery, it can also be used in the following scenarios:

[0055] 1) Industrial boiler; the tail gas generated by industrial boiler is usually high in pressure and temperature, and contains a large amount of energy.

[0056] 2) Used for thermal power generation device; thermal power generation will produce a large amount of high-temperature and high-pressure exhaust gas, and these exhaust gases are the energy source for the device to generate electricity.

[0057] 3) Used for transportation vehicles; vehicles such as cars and ships will continuously produce a large amount of high-temperature and high-pressure exhaust gas. The application of the device to vehicles can greatly reduce the loss of exhaust gas energy and improve the utilization efficiency of fossil energy.

[0058] In one specific embodiment, a liquid rocket engine test bench tail gas energy efficient utilization system is designed.

[0059] Firstly, the design condition of the turbine power generation device is determined to determine the design point of the turbine and the nozzle. As shown in Tables 1 and 2, the design condition of the device is assumed based on the energy utilization after the exhaust gas is blown off after the test. Since the device uses four adjustable nozzles, the flow rate of a single nozzle is 0.33 kg / s.

[0060] The design of the subsonic convergent section of the nozzle requires that the airflow be uniformly accelerated, the flow field at the throat section be uniform, and the flow not be separated. At the same time, the outer dimension is minimized to reduce the structural mass of the convergent section of the nozzle and reduce the friction loss. Therefore, the convergent section of the nozzle adopts a double-arc design.

[0061] The divergent section of the nozzle adopts the design method of a special-shaped nozzle. The generatrix of the special-shaped nozzle is a curve, and from the geometric shape and the processing angle, the simplest curve is a circular arc. When the nozzle expansion is small, the double-arc special-shaped nozzle is widely used, and here a double-arc nozzle is adopted.

[0062] Therefore, according to the upstream flow parameters and the nozzle design criteria, the technical indicators of the nozzle, i.e., the design parameters, are obtained as shown in Table 3. The sharp cone needle cone can produce fast response, and the smaller the surface taper, the stronger the area regulation ability, and the faster the system response.

[0063] Since the expansion of the gas only occurs in the nozzle (i.e., the nozzle) assembly, the flow process of the working medium in the adjustable throat area nozzle assembly is regarded as an isentropic process, and the isentropic expansion work of the working medium in the adjustable throat area nozzle assembly satisfies the equation:

[0064]

[0065] wherein, k is the specific heat ratio of the turbine working medium, R is the universal gas constant, is the total temperature at the inlet of the turbine, π is the isentropic expansion work of the gas in the nozzle.

[0066] According to the inflow conditions in Table 1, the specific heat ratio k and the total temperature at the inlet of the nozzle (i.e., the total temperature at the inlet of the turbine) can be determined; according to the turbine device expansion ratio (i.e., the turbine pressure ratio) in Table 1, the adjustable throat area nozzle assembly inlet and outlet pressure ratio can be determined.

[0067] Then, the actual speed of the adjustable throat area nozzle assembly outlet is:

[0068]

[0069] wherein can be taken as 0.92-0.95;

[0070] but, satisfy:

[0071]

[0072] Introducing the concept of circumference efficiency, circumference efficiency That is, the work done per revolution when a unit flow passes through a certain stage. Its ideal available energy in this level The ratio of the two parameters has the greatest impact on the circumferential efficiency of the impeller cascade. With the actual velocity at the nozzle exit ratio .

[0073] For impulse turbines, when hour( Let be the nozzle outlet airflow angle, and take the extreme value of the circumferential efficiency. Then, the tangential velocity of the airflow at the average diameter is... satisfy:

[0074]

[0075] in, This is the nozzle exit angle, typically 20-25 degrees.

[0076] Based on the above calculations, the average diameter of the turbine rotor can be obtained as follows:

[0077]

[0078] Rotor working blade inlet height:

[0079]

[0080] in, and These are the upper and lower coverage areas of the blades, respectively. The nozzle height is the diameter of the circular nozzle. The blade cover value can be found in the literature. For impulse turbines, , .

[0081] Working impeller airflow outlet angle:

[0082]

[0083] In the formula, For turbine working fluid flow rate, , The relative motion stagnation temperature at the inlet of the working impeller blade cascade. The total relative pressure at the impeller inlet. For the converging velocity at the outlet of the working impeller, the above parameters can be obtained based on the flow conditions in Table 1, combined with the nozzle structure parameters in Table 3, For the total pressure recovery coefficient of the nozzle, For the partial admission of the nozzle.

[0084] The inlet angle of the working impeller airflow:

[0085]

[0086] The width of the moving blade Related to its ability to resist bending and blade profile loss, generally, for longer blades For longer blades .

[0087] The chord length of the working blade:

[0088]

[0089] For the number of blades of the rotor, it can be calculated by the Zweifel coefficient. Zweifel (1945) established a calculation formula for calculating the optimal pitch / chord ratio, from which the cascade pitch can be obtained:

[0090]

[0091] wherein, is the cascade pitch, is the axial chord length of the blade, i.e. , , are the axial velocity components at the inlet and outlet of the channel, respectively. Zweifel found through research on a large number of turbine cascade experimental results that the optimal load coefficient is between 0.8 and 1.2. Therefore, the cascade pitch can be calculated, and the number of blades can be obtained.

[0092] Through the above design calculation, the average diameter of the turbine rotor, the chord length of the working blade, the blade setting angle, the number of rotor blades, and the inlet and outlet angles of the blade , and other turbine moving blade parameters such as Table 3 can be obtained.

[0093] That is, according to the flow rate, pressure ratio, inlet pressure, outlet pressure in Table 1, the contraction ratio, expansion ratio, cylinder segment diameter, throat diameter and nozzle exit diameter of the nozzle assembly shown in Table 2 can be calculated, and then based on the selected double-arc contraction segment design and special-shaped nozzle expansion segment design, the cylinder segment length and the remaining detailed structure size parameters can be obtained. Based on the impulse turbine assumption, the nozzle exit velocity can be obtained according to the inlet temperature and turbine pressure ratio parameters in Table 1, the turbine rotor average diameter can be obtained according to the nozzle exit angle and outlet velocity, the blade height can be calculated according to the nozzle diameter, the number of blades can be calculated according to the Zweifel coefficient, and the blade chord length can be calculated according to the calculated setting angle (i.e. the working impeller airflow exit angle, the working impeller airflow inlet angle). Finally, based on the nozzle exit angle and outlet velocity, rotation speed, rotor diameter and other parameters, combined with the turbine design criteria, the inlet and outlet angles of the blade can be calculated.

[0094] In summary, for the relationship between the throat area and the turbine design, the throat area of the nozzle without a needle cone and the nozzle exit velocity are determined according to the design condition expansion ratio and flow rate, and some design parameters of the turbine can be determined according to the nozzle exit velocity and nozzle exit angle. When the working condition changes, the outlet velocity can be adjusted by adjusting the throat area, so that the turbine rotor works at the design point (including: flow rate, rotation speed, outlet pressure, outlet temperature, calculated nozzle exit temperature and outlet pressure).

[0095] For the adjustable turbine power generation device, first, the nozzle inlet airflow parameters are sensed by pressure and temperature sensors, the control system adopts the PI algorithm commonly used on the adjustable Venturi tube of liquid rocket engines, and a PLC controller is used to control the needle cone.

[0096] Table 1 Design condition of turbine power generation device

[0097]

[0098] Table 2 Turbine device structure design parameters

[0099]

[0100] Table 3 Nozzle structure design parameters

[0101]

[0102] Table 4 Turbine moving blade parameters

[0103]

[0104] Any combination of the technical features in the above embodiments can be made, and for the sake of brevity, not all possible combinations are described above, however, as long as the combination of the technical features does not exist in contradiction, it shall be considered within the scope of the present disclosure.

[0105] The above embodiments only express several implementation manners of the present application, and the description is relatively specific and detailed, but it shall not be understood as a limitation on the patent scope of the present application. It shall be pointed out that, for ordinary skilled persons in the art, several modifications and improvements can be made without departing from the concept of the present application, and these shall be within the protection scope of the present application. Therefore, the protection scope of the patent of the present application shall be subject to the appended claims.

Claims

1. A liquid rocket engine test stand exhaust energy efficient utilization system, characterized in that, The utility model relates to a kind of turbine generator device, comprising: air source, adjustable throat area nozzle assembly, turbine generator device and controller; The air source is connected with one end of the adjustable throat area nozzle assembly by first pipeline, and the other end of the adjustable throat area nozzle assembly is connected with the turbine generator device by second pipeline; The controller is connected with the first pipeline, the second pipeline and the adjustable throat area nozzle assembly respectively, to control the throat area change of the adjustable throat area nozzle assembly according to the monitored pipeline parameter. The adjustable throat area nozzle assembly comprises motor, needle cone and nozzle; 2. The liquid rocket engine test stand exhaust energy efficient utilization system of claim 1, wherein, The motor is detachably connected with one end of the needle cone; The central axis of the needle cone is collinear with the throat center of the nozzle, and the other end of the needle cone has a gap with the throat of the nozzle; The controller is connected with the motor, and takes the pipeline parameter as input, outputs control signal, moves the needle cone in the direction of nozzle throat or away from nozzle throat by controlling motor, to realize the throat area change of the adjustable throat area nozzle assembly. The nozzle comprises convergent section, throat and divergent section connected in sequence.

3. The liquid rocket engine test stand exhaust energy efficient utilization system of claim 1 or 2, wherein The convergent section adopts double-arc design.

4. The liquid rocket engine test stand exhaust energy efficient utilization system of claim 3, wherein, The divergent section adopts special nozzle design.

5. The liquid rocket engine test stand exhaust energy efficient utilization system of claim 4, wherein, Temperature sensor and pressure sensor are arranged on the first pipeline; 6. The liquid rocket engine test stand exhaust energy efficient utilization system of claim 1 or 2, wherein The temperature sensor and the pressure sensor are connected with the controller, so that the controller controls the throat of the nozzle according to the temperature monitored by the temperature sensor and the pressure monitored by the pressure sensor. The turbine generator device comprises turbine moving blade, transmission mechanism and generator; 7. The liquid rocket engine test stand exhaust energy efficient utilization system of claim 1 or 2, wherein One end of the turbine moving blade is connected with the nozzle, and the other end is connected with the generator through the transmission mechanism. The turbine moving blade adopts single-stage impulse turbine.

8. The liquid rocket engine test stand exhaust energy high efficiency utilization system of claim 7, wherein, The isentropic expansion work of working medium in single-stage turbine satisfies the equation:

9. The liquid rocket engine test stand exhaust energy high efficiency utilization system of claim 8, wherein, The tangential velocity of airflow at the average diameter of the turbine moving blade satisfies: wherein L ad is the isentropic expansion work, Ω is the turbine reaction, R is the universal gas constant, k is the turbine working fluid specific heat ratio, is the turbine inlet total temperature, π is the turbine pressure ratio isentropic expansion work done by the gas in the nozzle.

10. The liquid rocket engine test stand exhaust energy high efficiency utilization system of claim 9, wherein, ​ wherein u is the tangential velocity of the gas stream at the average diameter, is the actual velocity of the nozzle exit, α 1 is the nozzle exit angle.

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