237 results about "Turbojet engine" patented technology

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 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 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.

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 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.

The invention relates to the technical field of engine recovery, in particular to a turbojet engine ejection recovery system and method in an aircraft fault state, and the system comprises an ejection mechanism which is arranged between a turbojet engine and an aircraft and used for disconnecting the turbojet engine from the aircraft. The turbojet engine is ejected out of the aircraft; the cut-off mechanism is arranged in the aircraft and used for cutting off an oil way and a circuit of the turbojet engine. The protection mechanism comprises a parachute assembly and an air bag assembly, the parachute assembly is connected with the turbojet engine, and a parachute is automatically opened in the landing process of the turbojet engine; the air bag assembly is folded and contracted outside the turbojet engine and automatically expands to wrap the turbojet engine in the landing process. The recycling device is compact in structure and flexible and convenient to use, recycling of the turbojet engine can be efficiently and safely completed when the fixedly-mounted turbojet engine is in an emergency, and the technical requirement for recycling of the turbojet engine in the field is met.

The invention belongs to the technical field of small turbojet engines, and particularly relates to a temperature field and pressure field prediction method and system for a gas compressor of a small turbojet engine. The method comprises the following steps: S1, acquiring data of different working conditions and data corresponding to a temperature field and a pressure field, enabling an acquired time sequence to correspond to the data of the temperature field and the pressure field, and then dividing into a training set and a test set; s2, establishing a neural network prediction model based on a gating circulation unit for the temperature field and the pressure field; s3, training a neural network prediction model based on a gating circulation unit for the temperature field and the pressure field; s4, testing the effect of the neural network prediction model based on the gating circulation unit for the temperature field and the pressure field; and S5, realizing temperature field and pressure field prediction of the small turbojet engine compressor. According to the invention, the problem that the neural network is affected by accumulative errors due to gradient disappearance and long prediction time in the unsteady flow field application is solved.

The present application relates to turbojet engine technical field, disclose a kind of combustion chamber structure and turbojet engine, combustion chamber structure includes: casing, flame tube and hydrogen nozzle;Flame tube is arranged in the inside of casing, the airflow passage between its outer wall and the inner wall of casing is formed, for cooling the outer surface of flame tube and providing the required air for combustion process.Hydrogen nozzle is made of built-in main pipe and nozzle piece, built-in main pipe is located in airflow passage, by with external hydrogen supply pipeline connection, realize the efficient delivery of hydrogen.Nozzle piece one end is connected built-in main pipe, the other end is worn in the side wall of flame tube and extends to the inside of flame tube, ensure that hydrogen can be accurately sprayed into combustion zone.The design greatly optimizes combustion efficiency, while significantly simplifying combustion chamber structure.

The utility model provides a three-dimensional vector thrust testboard for a micro turbojet engine, and relates to the technical field of engine thrust measurement. The test board comprises a rack fixed part and a rack movable part. The front end of the rack movable part is movably connected with the front support, and the rear end of the rack movable part is erected on the rear support; a tested engine is connected to the rear end of the movable part of the rack through a fixed clamping ring; an air inlet channel of the engine faces the rack movable part, and the central axis of the engine is located in the horizontal plane of the rack movable part. An axial thrust meter for measuring the axial thrust of the engine in real time is arranged between the front support and the front end of the movable part of the rack; a transverse thrust meter is arranged between each of the two sides of the rear end of the movable part of the rack and the rear support; and a vertical thrust meter is arranged between the rear support and the fixed clamping ring. In this way, the engine thrust measurement precision is improved, the influence of the structure on force measurement is reduced, and it is guaranteed that data measurement is more stable.

The invention relates to a whole machine dynamic balance system for a turbojet engine. The detection tool of the dynamic balance system is installed on the top face of a base of a conveying assembly, the number of the roller sets is multiple, each roller set comprises two rollers which are sequentially arranged in the first direction, the rollers are rotationally connected to the bottom face of the base through a roller frame, and a through locking hole is formed in the top of the roller frame; the two pushing pieces are oppositely arranged on the top face of the base in the first direction. The connecting rod is movably connected with the top face of the base in the first direction, and the two ends of the connecting rod are connected with the two pushing pieces through locking mechanisms correspondingly. The two ends of a connecting frame of the locking mechanism are connected with a pushing piece and a connecting rod, a driving pin is connected with the connecting frame, at least part of the driving pin is arranged in the connecting frame, a locking block is arranged in the connecting frame in a sliding mode in the first direction, the side, facing the driving pin, of the locking block is an inclined face, and a locking rod is connected with the top of the locking block and inserted into a locking hole. According to the engine base, the phenomenon that the engine on the base moves and collides with other objects can be avoided.

The utility model discloses a low temperature nozzle belongs to turbojet engine technical field, including the outside shield of fixed connection in the engine tail nozzle, the side of outside shield near engine body is equipped with a plurality of air inlet at interval, and the one end of outside shield near engine tail nozzle spout is equipped with the air outlet. When engine body runs, the spout of engine tail nozzle sprays high temperature high -speed gas, at this moment, utilize siphon principle, will bring out the airflow in outside shield, and cold air keeps on continuously acting between engine tail nozzle and outside shield, and the cooling of engine tail nozzle is given, and outside shield also will break off the heat radiation that inside tail nozzle emits, reaches the effect that reduces engine tail nozzle temperature.

1. Name of the product in this design: Electronic Controller for Turbojet Engine (ECU). 2. Purpose of this design: To control engine starting, acceleration, deceleration, steady state and stopping according to input control commands, as well as to check the function of engine accessories (starter motor, oil pump, solenoid valve, igniter, etc.), store engine data, and cool the engine. 3. The key design features of this product are its overall shape. 4. The image or photograph that best illustrates the design's key features: the front view. 5. This design is a thin product, so the left view is omitted; this design is a thin product, so the right view is omitted; this design is a thin product, so the top view is omitted; this design is a thin product, so the bottom view is omitted; this design is a thin product, so the perspective view is omitted.

The application discloses a turbojet engine detection ECU circuit board performance state device, including: a detection module connected with the ECU circuit board, used for detecting the performance state of the ECU circuit board, obtaining a detection result; a display module connected with the detection module, used for displaying the detection result; wherein the detection module includes an oil pump, an electromagnetic valve, a hot fire head, a starting motor and a temperature sensing probe. The device can effectively detect the performance state of the ECU circuit board, improve the detection efficiency and precision, reduce the cost, and provide valuable performance data for manufacturers.

PROPULSION ASSEMBLY FOR AN AIRCRAFT COMPRISING A TURBOJET ENGINE, A MAST AND MEANS FOR ATTACHING THE TURBOJET ENGINE TO THE MAST The invention relates to a propulsion assembly (100) of an aircraft having a vertical longitudinal axis (X) and median plane (P) and comprising a front casing (112a) with a rear face (113) perpendicular to the longitudinal axis (X), a primary structure (106) with a front wall (106e), and an engine attachment (150) comprising a female clevis (152a) integral with the front wall (106e), a male clevis (152b) integral with the rear face (113) and articulated in the female clevis (152a), a pin (154a) integral with the primary structure (106) and whose axis is parallel to the longitudinal axis (X) and a housing (154b) of the rear face (113) in which the pin (154a) is housed, and where the clevises (152a-b), the pin (154a) and the housing (154b) are arranged at the level of the median plane (P). Fig. 3

The present invention discloses a turbine afterburner matching system for a small turbojet engine supercharged afterburner combustion chamber. The system comprises a mixing channel and a supercharged annular cavity. The mixing channel communicates with the supercharged annular cavity, and an anti-backflow unit is provided between the mixing channel and the supercharged annular cavity. The anti-backflow unit is used to control the directional flow of fluid through the mixing channel into the supercharged annular cavity. This ensures a stable front-to-rear pressure drop ratio for the turbine in the supercharged afterburner combustion chamber of the small turbojet engine, prevents premature fuel autoignition, and ensures the stability of the turbine's operating state.

The present invention discloses a control method and control system for a micro-turbine engine, which primarily includes automatic correction of the starting fuel quantity and fuel closing control. At the beginning of engine startup, an initial starting fuel quantity is given and an initial fuel supply PWM signal is output. A determination is made as to whether the speed increase rate dn is greater than a preset rate of change D. If dn is greater than D, the starting fuel quantity is decreased by a preset fuel reduction amount and a corrected fuel supply PWM signal is output. If dn is not greater than D, the starting fuel quantity remains unchanged and a current fuel supply PWM signal is output. A determination is made as to whether the fuel growth rate is less than the speed growth rate. If so, the current fuel supply PWM signal remains unchanged. If not, the fuel growth rate is equal to the limited fuel growth rate, and a limited fuel supply PWM signal is output. This method enables the micro-turbojet engine to start quickly and enter an idling state quickly, improving the start success rate. During the throttle application process from idling to rated speed, the speed response is fast and speed fluctuation is minimal.

The invention relates to a hydraulic control circuit (11) for an orienting actuator (14) for orienting fan blades of a turbojet engine, the circuit (11) comprising an oil reservoir (17) and an axial piston pump (18) comprising a tiltable plate (23), as well as a control chamber (24) and a compensation chamber (27) for increasing the tilt of the plate (23) when the pressure inside these chambers increases, the pump (18) being supplied by the reservoir (17) and having its outlet configured so as to be able to be connected by a supply line (19) to one of the chambers (12, 13) of the orienting actuator (14), the other chamber of the orienting actuator (14) being connected to the reservoir (17), this control circuit comprising a valve (42) controlled by a computer (FADEC), which valve (42) allows the outlet of the pump (18) to be connected to the reservoir (17) in order to deactivate this pump (18).

The invention discloses a crawler-type remote control camouflage protection smoke screen vehicle and a working method, the crawler-type remote control camouflage protection smoke screen vehicle comprises a chassis, a smoke generating device and a box body, the chassis is a crawler-type mobile vehicle, the smoke generating device and the box body are both arranged on the chassis, and the smoke generating device is arranged in front of the box body; the fuming device is installed on the chassis through a left-right rotating head driving device and an up-down rotating head driving device, a turbojet engine and a spraying pipe are arranged in the fuming device, and the spraying pipe is arranged at the front end of the turbojet engine; a left-right rotating head limiting switch is arranged on the left-right rotating head driving device, and an up-down rotating head limiting switch is arranged on the up-down rotating head driving device; a smoke master control system is arranged in the box body, and the smoke generating device, the moving mechanism, the left-right rotating head driving device and the up-down rotating head driving device are all connected with the smoke master control system. The problems that an existing smoke screen vehicle is low in rapid transfer capacity on complex terrains and cannot rapidly adapt to environmental changes are solved.

The invention relates to a turbojet engine afterbody comprising:-an exhaust casing (9) carrying a manifold plate (11) surrounding a primary air flow and surrounded by a secondary air flow; -an afterburner (12) comprising a casing (13) and a bushing (16) delimiting together with the casing (13) an annular space (17) extending along the axis (AX), the diameter of the bushing (16) being greater than the diameter of the converging plate (11); -a fixed shroud (22) extending from the bus plate (11) to the inner liner (16), the fixed shroud (22) comprising a window (27); -a moving cover (23) extending along the fixed cover (22) and comprising a window (28), the cover (23) being rotatably movable about an axis (AX) between an open position, in which the window (27) in the fixed cover is open, and a closed position, in which the window (27) in the fixed cover (22) is closed by the moving cover (23).

The utility model discloses a heat dissipation structure of a smoke generator of a turbojet engine, which belongs to the technical field of heat dissipation structures and comprises a heat dissipation component, the heat dissipation component comprises a first heat dissipation sleeve and a second heat dissipation sleeve, and the first heat dissipation sleeve and the second heat dissipation sleeve are butted and then wrapped outside a tail spray pipe. Positioning support lugs are arranged on the sides, close to each other, of the first heat dissipation sleeve and the second heat dissipation sleeve; and the liquid flow channel is formed in the first heat dissipation sleeve and the second heat dissipation sleeve, and after the first heat dissipation sleeve and the second heat dissipation sleeve are in sealed butt joint, the liquid flow channel forms a through type net-shaped heat dissipation flow channel. Through the arrangement of a butt-joint structure, the first heat dissipation sleeve and the second heat dissipation sleeve are wrapped on the outer wall of the tail spray pipe, internal liquid circulates through communication of internal heat dissipation flow channels of the first heat dissipation sleeve and the second heat dissipation sleeve, efficient heat dissipation treatment of the tail spray pipe is achieved, and therefore the service life of all parts of the smoke generator is prolonged.

The invention discloses a method for obtaining a control strategy of a turbojet engine and a method for controlling the turbojet engine, the method for obtaining the control strategy of the turbojet engine comprises at least two iterations, and each iteration sequentially comprises a measurement and control stage and an updating stage; the measurement and control stage comprises at least two times of measurement and control and an array set is finally obtained; in the updating stage, the to-be-updated control strategy is updated through a primary function approximation method based on an array set to form a control strategy of the iteration and a control parameter used for representing a mapping rule from an independent variable to a dependent variable in the control strategy of the iteration; the method for controlling the turbojet engine is based on the obtained control strategy, the blade rotating speed is measured every set period, the fuel flow speed is obtained through calculation based on the control strategy, and then the turbojet engine is controlled according to the fuel flow speed obtained through calculation. By the adoption of the technical scheme, under the condition that dynamic parameters in the turbojet engine are unknown, continuous control over the fuel flow speed can be achieved, and the blade rotating speed reaches the target blade rotating speed.

The invention relates to the technical field of micro turbojet engine control, and discloses a control method for solving heat dissipation oil injection recombustion of a micro turbojet engine, which comprises the following steps of: closing fuel supply, establishing rotating speed closed-loop control, and keeping the actual rotating speed of the engine equal to a preset cooling rotating speed under disturbance; under the working condition, the actual cooling rate is obtained, and the reference cooling rate is obtained by inquiring the reference cooling model; comparing the actual cooling rate with a reference cooling rate to determine a rate deviation; according to the method, active rotating speed closed-loop control is established, a constant diagnosis reference working condition is forcibly constructed, and catastrophic interference of external aerodynamic disturbance on diagnosis logic is eliminated, so that the electronic control unit can still reliably sense and actively suppress hidden reburning under the working conditions of shutdown gliding and the like.

AIRCRAFT PROPULSION SYSTEM COMPRISING A TURBOJET ENGINE, A MOUNTING MAST AND MEANS FOR ATTACHING THE TURBOJET ENGINE TO THE MOUNTING MAST The invention relates to an aircraft propulsion system (100) comprising a turbojet engine (102) having a front casing (103) with a rear face (103a) and a rear casing (105), a mooring mast (104) with a front part (106a) and a lower spar to which a rear fitting (108) is attached, a front engine mount (150) fixed between the front casing (103) and the front part (106a), an arch (152) with a base (152a) fixed to the rear fitting (108) and two arms (152b-c), each having a first end integral with the base (152a) and a second end fixed to the rear casing (105), two upper bars (155) hinged to the rear face (103a) and embedded in the arch (152), and two lower bars (157) hinged to the rear face (103a) and to the arch (152). Fig.2.