381 results about "Turbojet engine" patented technology

An aircraft propulsion system comprising a turbojet having a fan casing with a rear face, a pylon with a front part and a lower spar with a rear fitting fixed to it, and a front engine attachment fixed between an upper part of the fan casing and the front part, a front arch having two arms, each having a first end fixed to the lower spar and a second end fastened in an articulated manner to the rear face, and two lateral rods, each having a first end fastened in an articulated manner to the rear fitting and a second end fastened in an articulated manner to the rear face.

The invention relates to a rapid prediction method for a temperature field and a stress field of a dangerous part of a key component of a small turbojet engine, belongs to the field of optimization design of small turbojet engines, and solves the problems of high cost and low efficiency of simulating the temperature field and the stress field of the turbojet engine by an existing conventional numerical simulation method. Comprising the steps of establishing a turbojet engine key component geometric model; generating temperature field and stress field data sets under different working conditions; reducing the order by using a POD flow field order reduction model and obtaining a transformation relation; training the deep learning model to obtain a mapping relation; inputting to-be-calculated working condition data to obtain reduced-order data; the distribution of the temperature field and the stress field under the specific working condition is obtained through recalculation by means of the transformation relation, rapid prediction can be achieved, and cost is saved. According to the method, deep learning and a flow field reduced order model are adopted, local dangerous parts under the given working condition can be predicted in time, calculation and time cost is saved, and timely data feedback is provided for prolonging the service life of an engine and keeping good performance of the engine.

The invention provides an aircraft framework and an aircraft, the aircraft framework is used for the aircraft flying in the atmosphere, the aircraft framework comprises a cage serving as a framework main body, the cage bears the self weight and the flight load of the whole aircraft, at least two first turbojet engines can be contained in the cage, and the first turbojet engines are arranged in the cage; the cage is open at least in the lower portion and used for allowing the jet flow of the first turbojet engine to pass through. Through the arrangement of the cage which is a skeleton main body, a power source device of the aircraft is collected and gathered, so that the miniaturization of the whole aircraft is realized.

The invention provides a personal aircraft chassis and a personal aircraft. The chassis comprises a main framework and a chassis body, the personal aircraft comprises a main framework and a thrust array assembly, the thrust array assembly is fixed to the main framework, thrust of the thrust array assembly is achieved through a plurality of first turbojet engines, the geometric position relation between the thrust action point and the thrust direction of the thrust array assembly and the main framework is kept fixed, and when the personal aircraft stops on the ground, the thrust array assembly is fixed to the main framework. The thrust direction of the thrust array assembly is perpendicular to the ground. The vector nozzle assemblies are fixed to the main framework, thrust of the vector nozzle assemblies is achieved through the second turbojet engines, the central axes of the second turbojet engines are perpendicular to the central axis of the first turbojet engine, and vector nozzles of the vector nozzle assemblies at least have one degree of freedom relative to the main framework. And high-dynamic, high-fault-tolerance and reconfigurable flight control which cannot be realized by utilizing pure thrust difference control can be realized.

The invention provides a fuel injection device capable of fully mixing air and hydrogen fuel, the combustion stability and the combustion efficiency of the hydrogen fuel are improved, the problem of local overheating of a hydrogen fuel combustion chamber is avoided, and the pure hydrogen fuel injection device of the micro turbojet engine comprises an annular nozzle and a hydrogen fuel supply pipe which are arranged in the combustion chamber; the annular nozzle comprises a plurality of micro-mixing units, and each micro-mixing unit comprises an air channel and a fuel injection hole which are perpendicular to each other; a fuel gas road network is arranged in the annular nozzle and comprises a first gas path channel, a second gas path channel and gas holes; the first air channel and the second air channel are arranged in a front-back staggered mode in the axial direction of the annular nozzle, and the first air channel and the second air channel are communicated through the air holes. Each micro-mixing unit is communicated with the second gas path channel through a fuel injection hole, and the first gas path channel is communicated with the hydrogen fuel supply pipe.

The high-temperature-rise combustion chamber comprises a sudden expansion diffuser, a cartridge receiver, a combined air atomizing nozzle and a flame tube and can solve the problems that in the prior art, a high-temperature-rise combustion chamber of a small and medium-sized turbojet engine is low in fuel oil supply pressure and fuel oil flow is increased, so that fuel oil atomization is poor, the combustion efficiency is low, and the combustion efficiency is low. And the service life of the flame tube is difficult to meet the requirement due to great reduction of the cooling air quantity and increase of the cooling air temperature. An overall layout form of fuel oil radial grading and axial grading is innovatively adopted, and swirlers used for assisting fuel oil atomization all adopt a beveling hole mode; space compound angles are innovatively adopted for cooling holes in the flame tube, and the overall layout form of different sizes and different hole pitches is adopted in different areas. Fuel oil atomization quality and efficient combustion under various working conditions can be improved, the service life of the flame tube is prolonged, and meanwhile the machining process of the swirler is simplified due to the fact that the swirler adopts an inclined hole mode.

The invention discloses a bimorphic high-speed unmanned aerial vehicle capable of taking off and landing on the water surface and flying in ground effect and a control method, and belongs to the technical field of unmanned aerial vehicles, the bimorphic high-speed unmanned aerial vehicle comprises a vehicle body and a control device, the vehicle body is provided with an adjustable turbine jet driving mechanism and telescopic folding wings, and the tail of the vehicle body is provided with an empennage and a tail turbine jet driving mechanism. Meanwhile, the invention discloses a control method based on the two-form aircraft capable of taking off and landing on the water surface and flying in the ground effect, and the two-form high-speed unmanned aerial vehicle capable of taking off and landing on the water surface and flying in the ground effect and the control method are adopted. By adjusting the inclination angle, the distance and the height of the two front turbojet engines, operation and conversion of various forms of water sailing, ground running, ground effect flight and height flight are achieved, the use scenes of the unmanned aerial vehicle are greatly enriched, and the requirement for high-speed flight is met by adopting the telescopic folding wings.

A low-altitude single-person aircraft powered by a micro-turbojet engine relates to the field of manned aircraft. The invention solves the problems of existing jet aircraft suits, such as heavy weight, inconvenient posture adjustment, and easy injury to the pilot. The invention comprises a flight backpack and two arm propulsion devices (7); the flight backpack comprises: a safety belt (1), a shell (2), an oil supply system (3), an electrical system (4), and a backpack micro-turbine engine (5); the oil supply system (3) is installed inside the shell (2), the electrical system (4) is installed on the upper end surface of the shell (2), and the backpack micro-turbine engine (5) is installed on the lower part of the shell (2); wherein the two arm propulsion devices (7) are respectively installed on the left and right sides of the shell (2) of the flight backpack. The flight backpack of the invention is the most complex component of a low-altitude single-person aircraft powered by a micro-turbojet engine. The invention is used in aircraft.

The invention concerns an aircraft propulsion system (100) comprising a turbojet (102) having a fan casing (103) with a rear face (103a), a pylon (104) with a front part (106a) and a lower spar (106b) with a rear fitting (108) fixed to it, and a front engine attachment (150) fixed between an upper part of the fan casing (103) and the front part (106a), a front arch (152) having two arms (152b-c), each having a first end fixed to the lower spar (106b) and a second end fastened in an articulated manner to the rear face (103a), and two lateral rods (154), each having a first end fastened in an articulated manner to the rear fitting (108) and a second end fastened in an articulated manner to the rear face (103a). With such an arrangement, the forces are transmitted to the rear of the pylon.

The invention belongs to the technical field of main shaft measurement, and discloses an aviation turbojet engine main shaft measuring device and a use method thereof. The measuring device for the main shaft of the aviation turbojet engine comprises a mounting rack, two annular plates are symmetrically and fixedly connected to the mounting rack, an annular plate is movably connected between the two annular plates, a plurality of measuring assemblies are uniformly distributed on the annular plate, and a detection rotating assembly is arranged on the mounting rack. The power output end of the detection rotating assembly is in transmission connection with the annular plate and used for driving the annular plate to rotate, a movable plate is movably connected to the inner bottom face of the mounting frame, and two main shaft locking assemblies are movably arranged on the top face of the movable plate. A transverse displacement assembly used for driving the movable plate to move horizontally and transversely is arranged on the inner bottom face of the mounting frame. According to the invention, the detection and measurement operation of the main shaft can be efficiently and accurately completed.

The utility model discloses a cleaning device for a bearing sleeve of a turbojet engine, which is characterized in that a first sealing element and a second sealing component block the bottom end of the bearing sleeve and an oil outlet between an inlet and an outlet of a lubricating oil circuit pipeline, so that a cleaning medium can only flow into the lubricating oil circuit pipeline from a first through hole in the second sealing component; the first through hole is communicated with the second through hole through a cleaning circulation pipeline so as to be communicated with the lubricating oil way pipeline to form a circulation loop for cleaning the lubricating oil way pipeline; the third sealing assembly blocks the top end of the bearing seat sleeve and the inner wall of the hollow cavity, it is ensured that a cleaning medium can only flow into the cooling gas path pipeline from a third through hole in the second sealing assembly and flow out of a gas inlet of the cooling gas path pipeline, and the third through hole communicates with the gas inlet of the cooling gas path pipeline through a cleaning circulation pipeline. And a circulation loop for cleaning the cooling air channel pipeline is formed by communicating with the cooling air channel pipeline.

The invention relates to a thrust measuring device for a vertical-position micro turbojet engine. The device is used for measuring the thrust when a micro turbojet engine with a vector nozzle is vertically mounted, and mainly comprises two parts, namely a static frame and a movable frame. The static frame is composed of adjustable horizontal bottom feet and a static support, and an installation flat plate is arranged on the top of the static frame. A rotating shaft is fixed on one side of the static frame mounting flat plate, and the movable frame can tilt around the rotating shaft. The engine is vertically installed on the sliding rail of the movable frame and connected with the tension and compression sensor. A pressure sensor is fixed to the position, opposite to the rotating shaft, of the static frame installation flat plate, and the movable frame is supported on the pressure sensor. Through the design, the weight of the movable frame is larger than the maximum thrust of the engine, and it is guaranteed that the movable frame does not tilt in the engine test process. The thrust in the vertical direction of the engine can be calculated according to the torque relation by using the pressure value measured by the pressure sensor. And the measured value of the tension and compression sensor is the horizontal thrust generated by the vector nozzle. The device is simple in structure, convenient to install and small in measurement error.

The turbojet engine comprises an engine shell, a starting motor, a gas compressor, a combustion chamber and a turbine are arranged in the engine shell, and the exhaust end of the combustion chamber communicates with the atmosphere through an exhaust nozzle; the compressor is provided with an air inlet and an air outlet; the air outlets are divided into a first air outlet and a second air outlet; the first gas outlet is formed in the engine shell and communicates with the gas inlet end of the combustion chamber, so that part of compressed gas is introduced into the combustion chamber to be mixed with fuel oil for combustion, the turbine is impacted, and the fuel gas is discharged into the atmosphere to generate thrust; and the second air outlet extends out of the engine shell and is communicated with the external atmosphere, so that the other part of compressed air is directly discharged into the external atmosphere, and the posture of the whole aircraft is adjusted. According to the invention, through the design of introducing the second air outlet, the posture is adjusted by utilizing compressed air led out from the air compressor, so that the air which is not utilized originally is fully utilized, and the overall energy utilization rate is improved.

The present invention relates to a micro-engine test bench. A roll and pitch attitude test bench for a micro-engine, comprising a micro-engine and a support base, is characterized in that it further comprises a pitch test bench, which is mounted on the support base through a pitch device for driving its pitching motion; the micro-engine is fixedly mounted on a roll table, and the roll table is further mounted on the pitch test bench through a roll device for driving its rolling motion; a vibration sensor, a rotational speed sensor, an inclination sensor, and a tail nozzle temperature sensor are provided on the roll table, and all of them send induction signals to an electrical control cabinet. At the same time, the electrical control cabinet is also electrically connected to the pitch device and the roll device for controlling the pitch device and the roll device. The present invention realizes the precise control of the pitch angle and roll angle of a turbojet engine, simulates the real attitude of the micro-turbojet engine in the air, and accurately obtains various technical parameters when the micro-turbojet engine rolls and pitches.

The invention provides a turbojet engine starting control method and system based on artificial intelligence. The method comprises the steps that the oil way pressure of an ignition oil way of an engine, the environment temperature of the environment where the engine is located and the internal temperature of the engine are obtained; under the condition that the oil way pressure meets a preset condition and the environment temperature is smaller than or equal to a first temperature threshold value, an igniter is controlled to be started, a gunpowder starter is controlled to be started at a first time point, and the first time point is a time point which is delayed by a preset duration of the time point for controlling the igniter to be started; and under the condition that the oil way pressure meets the preset condition and the environment temperature is larger than a first temperature threshold value, or the oil way pressure meets the preset condition and the internal temperature is larger than a second temperature threshold value, the igniter and the gunpowder starter are controlled to be started at the same time. According to the self-adaptive control method, the starting reliability of the turbojet engine in the low-temperature environment is effectively enhanced, and meanwhile the requirement for quick starting in the normal-temperature or heat engine state is met.

The invention discloses a turbojet power logistics unmanned aerial vehicle based on a fixed wing structure and a control method. The unmanned aerial vehicle comprises a carbon tube fuselage, cork wood wings, a full-motion horizontal empennage, double vertical empennages, a nose fairing, a storage bin, a turbojet engine, an undercarriage, a flight control system and an emergency parachute. According to the turbojet power logistics unmanned aerial vehicle based on the fixed wing structure and the control method, the fixed wing structure and the turbojet engine are organically combined, short-distance take-off within 50 meters can be achieved, compared with a traditional propeller type fixed wing unmanned aerial vehicle, the take-off distance is shortened by 40% or above, the cruising speed can reach 80 km / h-120 km / h, the unmanned aerial vehicle can be matched with a standard material box of 200 mm * 200 mm * 400 mm, and the unmanned aerial vehicle is suitable for large-scale popularization and application. The flight control system supports two modes of manual remote control and automatic driving, and is equipped with an emergency parachute, so that safe parachute recovery of the unmanned aerial vehicle in an emergency state is realized, and the problem that a traditional unmanned aerial vehicle has obvious limitation when meeting emergency material delivery requirements is effectively solved.

The invention discloses a forming welding device for manufacturing a turbojet engine combustion chamber, and belongs to the technical field of welding machining. Two mounting plates are vertically mounted on the top wall of the base in a sliding manner, and clamping holes are formed in the side walls of the two mounting plates; a bracket is fixedly mounted on the side wall of the mounting plate; a driving mechanism for driving the two mounting plates to move synchronously is arranged on the base; an annular guide rail is arranged on the base, a mounting frame is slidably arranged on the annular guide rail in a sleeving mode, and a welding machine is arranged on the mounting frame. Due to the fact that the temperature of a welding part rises in the welding process, in the process that a rotating pipe drives a first exhaust hole to rotate, the inner wall of the welding part can be impacted through airflow, at the moment, under the action of heat exchange, the airflow can be heated, and high-temperature airflow in a closed combustion chamber preheats a low-temperature area through heat exchange; the whole temperature uniformity of the wall surface of the combustion chamber is effectively improved in the process, so that the temperature gradient is obviously reduced during welding, the thermal shock is inhibited, and the thermal stress is reduced.

The invention relates to the technical field of unmanned aerial vehicles, in particular to a turbojet power system of a multi-rotor unmanned aerial vehicle, which comprises a turbojet engine, a fuel supply module, an electronic control unit, an air inlet flow guide structure, an exhaust flow guide structure, a cooling module, a damping bracket, a quick disassembly and assembly interface, a redundant backup system and an energy recovery module, according to the multi-rotor unmanned aerial vehicle, the power performance of the multi-rotor unmanned aerial vehicle is improved by replacing a traditional motor with the micro turbojet engine, so that the unmanned aerial vehicle can obtain longer endurance time under the condition of carrying the same load, and particularly, the unmanned aerial vehicle can still keep stable power output in extreme environments such as high altitude or high temperature; the problem of power attenuation caused by thin air or poor heat dissipation of the motor is solved, multi-degree-of-freedom thrust direction adjustment is achieved through the design of the deflectable nozzles, and the unmanned aerial vehicle can still keep accurate flight attitude control under severe environment conditions such as strong wind.

An acoustic treatment panel including acoustic absorption cells each comprising a bottom wall, a porous inlet wall, an enclosure extending in an axial direction between the bottom wall and the inlet wall, and a first acoustic horn extending inside the enclosure between a first opening and a second opening that is smaller than said first opening, the first opening facing said inlet wall. Each cell comprises a second horn extending inside the enclosure in the first axial direction between a first opening of the second horn and a second opening of the second horn that is smaller than said first opening.

A thrust reverser device for an aircraft turbojet engine nacelle an electro-hydrostatic actuation system that includes a motor pump and a first actuation assembly. The first actuation assembly includes a series of at least two actuators configured to be attached to a movable cover of the thrust reverser and to move the movable cover or covers between a retracted position and a deployed position relative to the nacelle. The first electro-hydraulic actuation system further includes a volumetric flow divider with an inlet connected to the motor pump and at least two outlets, wherein each outlet is connected to an actuator. The volumetric flow divider is configured to receive a fluid and to supply the fluid to the actuators with an equal flow rate.

The present invention belongs to the field of aeroengines, and specifically relates to a high-throughput counter-rotating turbojet engine. The engine comprises a casing, the casing including an intake casing and an exhaust casing, the casing internally provided with: an air intake system, a compressor system, a combustion chamber system, a turbine system, a tail jet system, and a transmission shaft system; the compressor system decelerates and pressurizes air flowing in through the air intake; the compressor system comprises a first rotor and a second rotor in counter-rotation; the first rotor comprises a plurality of first rotor fans, the second rotor comprises a plurality of second rotor fans, and the first and second rotor fans are spaced apart; the present invention increases air flow and enhances engine thrust by enabling the first and second rotors to rotate in counter-rotation, while ensuring a small weight and volume.

The utility model discloses a generator control device, a gas generator system and an aircraft, the generator control device comprises an automatic control unit, a sensing unit and a rectification unit, the generator is connected with a turbine of a turbojet engine through a main shaft, and the generator is connected with electric equipment. The automatic control unit is connected with an oil pump of the turbojet engine, electric equipment, the sensing unit and the rectifying unit, the sensing unit comprises a rotating speed sensor connected with a main shaft, the rectifying unit comprises a diode rectifier bridge and a buck-boost converter, the generator is connected with the diode rectifier bridge, and the diode rectifier bridge is connected with the buck-boost converter. And the buck-boost converter is connected with electric equipment. The power supply can stably output electric energy according to the requirements of electric equipment, ensures that the output voltage is stable, and meets the requirements of the electric equipment.

The present invention discloses a turbojet engine and flight equipment. The turbojet engine includes a compressor; a combustion chamber connected to the compressor; a guide, the guide including a guide body and a guide housing, the guide housing being sleeved on the outside of the guide body, and the end face of the guide body being provided with a circumferential positioning groove, the guide body being coaxially sleeved with the main shaft sleeve of the turbojet engine, and the main shaft sleeve being provided with a matching positioning groove capable of matching with the circumferential positioning groove; a circumferential positioning member, the two ends of the circumferential positioning member being respectively inserted and matched with the circumferential positioning groove and the matching positioning groove; an axial positioning member for axially positioning the guide; and a tail nozzle assembly for ejecting reverse thrust airflow. The present invention fixes the guide by means of circumferential positioning members and axial positioning members, thereby reducing the contact area between the guide and the main shaft sleeve, reducing the heat transferred to the main shaft, thereby reducing the operating temperature of the bearings on the main shaft and extending the service life of the bearings.

Hybrid turbomachine comprising: an input shaft (a26) configured to be driven by a low-pressure turbine (102); an output shaft (a64) to drive an electric machine (104) and another output shaft (a64') to drive another electric machine (104'); a speed variator (106) to adjust the speed of each output shaft (a64, a64') according to a speed of the input shaft, comprising: a first planetary gear train (T1) having, at the input, a planet carrier (108) connected to the input shaft, a ring gear (112) and a planet (110); at the output, a first sun gear (130) connected to said output shaft, a second planetary gear train (T2) having, at the input, the ring gear (112) and another planet (120); At the output, a second sun (140) is connected to the other output shaft, and a means (MB) is provided for locking or freeing the crown. Figure for the abbreviation: Fig. 1

This utility model discloses an adjustable two-dimensional tail nozzle, belonging to the field of two-dimensional tail nozzle technology; it includes a connected engine and an adapter casing. A servo is provided on one side of the adapter casing. The output side of the adapter casing consists of a rectangular opening formed by two oppositely arranged fixed plates and two oppositely arranged adjusting plates. A power arm is provided at the output end of the servo. The right end of the power arm is connected to two parallel first connecting rods by a pin shaft. The other ends of the two first connecting rods are respectively connected to two second connecting rods. It can be applied to small afterburning turbojet engines to improve the flight speed of UAVs. Its structure is simple. The synchronous opening and closing movement of the adjusting plates is achieved through a single servo drive source, reducing the control difficulty, realizing engine thrust adjustment, reducing the number of structures, and making it lightweight, highly reliable, and low in cost. By setting a support rod on the outside of the adjusting plate, the adjusting plate can be limited. The structure is simple and the use effect is good.

The embodiment of the invention relates to the field of removing ice and snow on railway turnouts, in particular to turbojet ice and snow removing equipment which comprises a body, a turbojet engine and an air compressor. A fuel inlet, an air inlet and a high-temperature and high-pressure gas outlet are formed in the body, the turbojet engine and the air compressor are arranged in the body, fuel enters the body from the fuel inlet so as to enter the turbojet engine, and external air enters the body from the air inlet so as to enter the air compressor; the turbojet engine generates high-temperature and high-pressure gas through fuel and external air, the high-temperature and high-pressure gas is output from the turbojet engine and exhausted from the high-temperature and high-pressure gas exhaust port, and the exhausted high-temperature and high-pressure gas purges the position to be purged. The body is configured to reduce vibration of the turbojet engine. According to the turbojet ice and snow removing equipment provided by the embodiment of the invention, accumulated snow, frozen snow and frozen ice on a railway are melted and blow-dried through high-temperature and high-pressure gas, ice and snow on the railway can be efficiently and comprehensively removed, and normal work of the railway is guaranteed.

A rear portion of a turbojet engine with converging-diverging nozzles includes a device for controlling the position of the diverging vane relative to the position of the converging vane, the device including a first link (120) having a first end (120A) hinged to the converging vane (30) and an opposite second end (120B), a second link (122) having a first end (122A) hinged to the diverging vane (32) and an opposite second end (122B), and a third link (124) having a first end (124A) hinged to a synchronization ring (82) and an opposite second end (124B). The second end (120B) of the first link is hinged to the second end (122B) of the second link and / or to the second end (124B) of the third link, and the second end (124B) of the third link (124) is hinged to the second end (122B) of the second link (122).

The invention relates to an aircraft turbine engine comprising at least one propeller shaft and an electric machine (1), the rotor (3) of which is rigidly connected to the propeller shaft, the electric machine (1) comprising at least three power channels (C1-C4), a control device (4) designed to determine a control parameter for each power channel (C1-C4) as a function of a control command (Com), and a correction device (5) designed to correct the control parameter for each power channel (C1-C4) as a function of the control command (Com). And a correction device (5) designed to correct a precession movement of the rotor (3) relative to the stator (2), the correction device (5) being designed to determine a control command (Com) of the control device (4) as a function of the setpoint command (Comcons) and the precession level (NP), the control command (Com) being designed to generate a total magnetic force (FG) in the opposite direction to the precession movement (P +) of the rotor (3) in order to achieve damping of the precession movement.

The application discloses a manual pipe bending device for a turbojet engine oil supply pipeline, which comprises a base plate, a positioning column, a rotating disc and a bending die; the top end of the positioning column is fixedly provided with a pipe bending module for pipeline bending; the rotating disc is slidingly sleeved outside the positioning column, a handle is fixedly connected to the sidewall of the rotating disc, the top end of the rotating disc is fixedly connected to a limiting block, and the limiting block is correspondingly arranged with the pipe bending module; an angle scale is arranged on the base plate, and the angle scale is used for determining the rotation angle of the rotating disc during pipe bending; the bending die is correspondingly arranged with the pipe bending module, and the bending die is used for determining the interval of pipeline bending. Through the arrangement of the pipe bending module, the pipeline is positioned, and through the bending die, the bending position and the bending interval of the pipeline are determined. When the pipeline is bent, the required interval does not need to be repeatedly measured, and after the pipeline is bent, the pipeline can be positioned again through the bending die and the pipe bending module, so that the operation is simple, and the consistency of batch production is high.

The utility model discloses a multi-model micro turbojet engine test bench, which relates to the technical field of engine testing and comprises a body, an oil tank is arranged on the inner bottom surface of the body and communicated with a plurality of connecting pipes, the tops of the connecting pipes extend out of the body, a turbojet engine is arranged on the top surface of the body and communicated with the connecting pipes through oil pipes, and the connecting pipes are communicated with the oil pipes. A control system is installed on the side face of the body and communicated with an external power source. The turbojet engine clamping device is simple in structure, convenient to use, capable of clamping turbojet engines of different models, wide in application range, reasonable in space application, convenient to operate and control, capable of achieving real-time observation and high in practicability.