233 results about "Turboshaft" patented technology

The embodiment of the invention provides a digital simulation modeling method and device of a tandem type hybrid electric propulsion system, computer equipment and a medium, and relates to the technical field of simulation modeling, and the method comprises the following steps: constructing a turboshaft prime mover model, constructing a generator model, connecting the turboshaft prime mover model and the generator model in series to generate a generator set, the power generation groups are connected in parallel to construct a turbine power generation system; constructing a propulsion system; constructing a battery energy storage system; constructing a power transmission and distribution system, and respectively connecting the power transmission and distribution system with a turbine power generation system, a propulsion system and a battery energy storage system; and constructing a comprehensive control and energy management system, and outputting a digital simulation modeling result through a turbine power generation system, a power transmission and distribution system and a battery energy storage system. According to the scheme, a simulation modeling method based on multi-physics field coupling double-power-source dynamic response and energy flow regulation and control provides reliable modeling technical support for engineering application of a tandem type hybrid electric propulsion system.

The invention discloses a turboshaft engine dynamic process performance parameter prediction method considering thermal inertia, and belongs to the field of aero-engine fault diagnosis. According to the method, the influence of thermal inertia on gas path parameters in the dynamic process of the turboshaft engine is analyzed, a thermal inertia coefficient is designed and introduced, a hot end component heat exchange model is established, a component characteristic diagram is corrected in combination with real test run data, and performance parameters in the dynamic process are predicted in real time through a neural network prediction algorithm. According to the method, the hysteresis phenomenon of gas temperature change in the dynamic process can be effectively described, the accuracy of dynamic modeling of the hot end component is improved, the capability of predicting dynamic thermodynamic parameters of the engine in real time is improved, and a scientific basis is provided for formulating a prediction health management strategy of the turboshaft engine.

The invention discloses a rapid dynamic response control method for a turboshaft engine. According to the method, a gas turbine rotation acceleration cascade control structure is used for controlling the turboshaft engine; a feed-forward compensation form with predicted required power as a feed-forward compensation amount is introduced into an inner ring fuel flow instruction control position of a gas turbine rotation acceleration cascade control structure; introducing a feed-forward compensation form which takes a required power prediction error as a feed-forward compensation amount into an outer ring gas turbine rotation acceleration instruction control position of the gas turbine rotation acceleration cascade control structure; and a recursive least square algorithm is used to carry out adaptive adjustment on feed-forward compensation parameters of the two loops. The invention further discloses a rapid dynamic response control system for the turboshaft engine. Compared with the prior art, the technical scheme can effectively improve the response speed of the engine control system.

The invention belongs to the technical field of turboshaft engines, and provides a fuel oil compensation method and device for starting ignition of a turboshaft engine and a storage medium, the method comprises the steps that whether ignition is successful or not in an ignition interval is detected, if ignition fails, corresponding fuel oil compensation is increased, and the ignition interval is that ignition is conducted within the preset range of the rotating speed of the engine; whether the turbine outlet temperature rise meets a first condition or not is detected, and if the first condition is met, fuel compensation is ended. The ignition success rate is improved by detecting whether ignition in the ignition interval succeeds or not and performing fuel compensation outside the designed ignition interval after ignition fails, so that the purpose of reducing the starting time is achieved, and in order to further improve the ignition success rate, whether the turbine outlet temperature rise meets the first condition or not is continuously confirmed, so that the starting time is shortened. And the ignition success rate is further improved, and good adaptability is achieved.

The invention discloses a method, a device and equipment for monitoring the degradation state of a turboshaft engine and a medium, and aims at overcoming the defects of an existing method for sensing the degradation state of the turboshaft engine based on limited measurement parameters. A degeneration factor range linear solving method based on a degeneration factor fingerprint map theory is introduced on the basis of a traditional particle swarm optimization algorithm, and a more accurate and reasonable particle optimization limitation range is provided for the particle swarm optimization algorithm under the condition that available system input parameters are insufficient. For any sensor actual acquisition main performance parameter degradation situation combination, a corresponding specific optimization range can be provided, and the situation that degradation factor optimization falls into local optimization in a plurality of possible optimization spaces due to insufficient data complexity is avoided. Therefore, degradation state monitoring of the turboshaft engine is optimized, and the overall performance degradation trend of the engine is accurately obtained.

The invention discloses a turboshaft engine component characteristic dynamic optimization method, device and equipment and a medium, and the method comprises the steps: S1, calculating to obtain typical parameters of an engine in each flight stage, after multiple rounds of typical parameter calculation results are obtained through iteration by changing the relative conversion rotating speed increment, the gas compressor characteristic diagram with the lowest fuel flow weighted average value is selected as the optimal optimization design result of the gas compressor; s2, performing engine parameter calculation according to the optimal optimization design result of the gas compressor to determine the position relationship between the highest efficiency area of the power turbine on the turbine characteristic diagram and each stage of the aircraft; the power turbine characteristic diagram is adjusted through a double-Gaussian correction formula, the power turbine efficiency of the design point is adjusted, then iteration is carried out, typical parameter calculation results of the engine in the flight stages of multiple rounds are obtained, and the power turbine characteristic diagram with the lowest fuel flow weighted average value is selected as the optimal optimization design result of the power turbine. The method is low in optimization cost, efficient and accurate.

The invention provides a generator and battery power distribution method and device based on model prediction control, and belongs to the technical field of power distribution. The method provided by the invention comprises the following steps: loading a preset dynamic model library; generating and calibrating a total load power prediction sequence based on the flight plan and the rotor aerodynamic power model; calling the dynamic model library to predict and control the rolling optimization to output the optimal power instruction sequence of the generator; in the rolling optimization execution process, bus voltage dynamic characteristics are monitored in real time, a graded response mechanism is started based on a turboshaft engine dynamic model when load sudden change occurs, and the power of a generator and the power of a battery are adjusted; in the load abrupt change response process, rolling optimization is executed again based on the adjusted power and the corresponding system state, meanwhile, feed-forward compensation is carried out on the bus voltage deviation, and the voltage is controlled to be within a stable interval set by rolling optimization by adjusting the power of the generator; and circularly executing until the flight task is finished, and determining the generator power and the battery power.

The invention provides a fuel-based, single-engine multirotor drone, wherein the drone comprises a body, a turboshaft engine to generate rotational movement, a plurality of rotors, a plurality of pulleys, a plurality of transmission belts, and a gearbox assembly. The gearbox assembly comprises a pair of coaxial counter-rotating gears which are coupled to a pair of coaxial counter-rotating shafts, wherein the rotational movement of the engine is transferred synchronously from the counter-rotating shafts of the gearbox to the plurality of the rotors through the plurality of pulleys and transmission belts.

The invention relates to a hybrid vertical take-off and landing aircraft energy management method based on a model predictive control algorithm, which belongs to the technical field of aircraft energy management, and comprises the following steps: establishing a size model and a performance model of a multi-rotor configuration based on a top demand, a hybrid architecture and a propulsion configuration by using an iterative scheme of size design optimization; forming constraint of size iteration, and optimizing required overall size parameters of the aircraft through an fmincon function; a component-level turboshaft engine model and a battery internal resistance model are provided, a theoretical model is established based on gas-heat dynamic characteristics of each component of the turboshaft engine, the battery internal resistance model is established based on battery testing, and a battery peak current is obtained; and an energy management strategy based on an MPC algorithm, DP optimization information, supplementary CS and a charge strategy.

The invention discloses an engine accessory on-condition maintenance decision-making method considering inherent characteristics and repair characteristics, and belongs to the field of turbine engine maintenance, and the method comprises the following steps: classifying turboshaft engine accessories based on the inherent characteristics to obtain classified accessories, the classified accessories comprise a repairable accessory and an unrepairable accessory; performing fault damage analysis on the classified accessories to obtain an external field maintenance decision; determining factory return maintenance accessories based on the external field maintenance decision; and analyzing the factory return maintenance accessory based on the repair characteristics to determine a maintenance decision of the accessory in the overhaul. The method can be used for guiding the maintenance decision of accessories of the turboshaft engine in external field use and factory return overhaul stages, scientifically determining the maintenance mode and the disassembly depth, improving the reliability and maintainability of the engine, reducing the maintenance cost and improving the guarantee capability of the engine.

The invention discloses a methanol / aviation kerosene dual-fuel chemical regenerative hybrid aero-engine, relates to the technical field of environment-friendly processing, and solves the problems that the electricity demand of turboshaft engine airborne equipment is increased and the carbon emission of the engine is too high. A core engine system comprises an aviation fuel storage tank, an air inlet channel, an air compressor, a combustion chamber, a turbine system, an exhaust nozzle and a rotor wing, wherein the aviation fuel storage tank is communicated with the combustion chamber; the air inlet channel, the air compressor and the combustion chamber communicate in sequence. One part of the methanol reforming product is introduced into a combustion chamber to be mixed with aviation fuel for combustion, and the other part of the methanol reforming product generates power through a fuel cell; the combustion chamber, the turbine system and the exhaust nozzle are communicated, and fuel gas generated by the combustion chamber is introduced into the turbine system to drive the turbine system; the turbine system drives the gas compressor and the rotor wing. The fuel cell is used for generating electricity to meet the electricity demand of airborne equipment, the shaft power on the engine main shaft is reduced, the engine output power is effectively improved, and the operation stability of a core engine system is improved.

The invention discloses a fault diagnosis system for an aviation turboshaft engine, and relates to the technical field of fault diagnosis. An acquisition module acquires horizontal, vertical and axial vibration signals; the processing module is used for generating real-time feature vectors by introducing dynamic short-time Fourier transform, variational mode decomposition and singular value decomposition of an unfixed-width Gaussian window; the incremental diagnosis module calculates the matching degree between a real-time feature vector or a prediction feature vector and a known class through a dynamic expansion network, selects to belong to the known class or triggers an expansion mechanism according to an adaptive resonance theory, constructs joint loss by using a Bayesian learning framework after expansion, and performs Gaussian process modeling; loss is determined through Bayesian optimization, and network parameters are updated; the prediction module generates a prediction feature vector when there is no fault; the maintenance module sends or stores a maintenance scheme according to needs, self-adaptive recognition of known classes and new fault classes is achieved, the state can be pre-judged in advance, the operation safety of the engine is effectively guaranteed, and the maintenance cost is reduced.

The invention belongs to the technical field of helicopter control, and provides a turboshaft engine identification and prediction control method based on MRR-KELM in order to solve the problem that a prediction model constructed for a turboshaft engine is poor in generalization, an MRR-KELM algorithm is combined with a Huber loss function and a self-adaptive regularization strategy, noise interference and abnormal value influence are restrained, and the prediction model is optimized. The prediction error of the constructed turboshaft engine prediction model is greatly reduced, and adaptive adjustment can be performed through Gaussian kernel dynamic mapping and working condition parameters; according to the algorithm, a regularization item and a robust loss function are introduced, the kernel function mapping capability is combined, the generalization performance and the calculation efficiency of the model under the complex working conditions of small samples, noise interference and the like are remarkably improved, further, the turboshaft engine prediction model is embedded into a nonlinear model prediction control framework, and the prediction accuracy of the turboshaft engine is improved. And high-precision stable control over the rotating speed of the power turbine is achieved through rolling optimization and feedback correction.

The utility model discloses a sleeve repair tool for a fuel manifold of a turboshaft engine, which comprises a support with the top end provided with a penetrating column with threads on the inner wall, and further comprises a threaded pressing rod with the top end provided with a pulling rod, and the threaded pressing rod is in threaded connection with the penetrating column through the threads on the outer wall; a semi-arc-shaped pressing plate is arranged at the bottom of the T-shaped sliding block, and the T-shaped sliding block is movably connected with the threaded pressing rod; a shape correcting arc plate is arranged on the positioning column opposite to the semi-arc pressing plate, and the positioning column is located at the middle section of the support; and the gland is connected to the positioning column through a bolt. According to the sleeve repairing tool for the turboshaft engine fuel main pipe, the problems that in the prior art, manual positioning is not accurate, and pipe wall cracks and internal threads are prone to being crushed to be scrapped due to repeated knocking are solved.

An aviation turbine shaft engine over-rotation protection execution mechanism detection device comprises an oil way system and a control circuit, and the oil way system comprises an oil tank, a direct-current electric pump and a two-position three-way electromagnetic valve; a normally open loop of the electric pump and the electromagnetic valve and the oil tank form a circulating oil path through an oil pipe; a normally closed loop of the electromagnetic valve is connected with an oil inlet of the aviation turboshaft engine over-rotation protection execution mechanism test piece, and an oil outlet of the over-rotation protection execution mechanism test piece is connected with the oil tank through an oil pipe; the control circuit comprises a power switch, a detection switch and a two-position three-way electromagnetic valve; the power switch controls power supply of the direct-current electric pump, and the detection switch controls power supply of a control coil of the two-position three-way electromagnetic valve and a test piece of the aviation turbine shaft engine over-rotation protection executing mechanism. According to the utility model, an integrated box type design is adopted, the size is small, the structure is simple, an onboard direct current 27V power supply is used for supplying power, the external field use requirement of a helicopter can be met, and the working condition of an aviation turbine shaft engine over-rotation protection executing mechanism can be conveniently detected.

The invention discloses a hybrid power turboshaft engine and a working method, the engine comprises a core machine system and a fuel cell power supply system, and the core machine system comprises an air inlet channel, an air compressor, a main combustion chamber, a gas turbine and a power turbine; the fuel cell power supply system comprises a liquid ammonia storage tank, a fuel pump, a liquid ammonia decomposer, a pressure reducing valve, a heat exchanger and a proton exchange membrane fuel cell. According to the invention, a hybrid power form of combining a turbine power system and a fuel cell power system is adopted, so that the output power of the turboshaft engine is ensured, and meanwhile, the proton exchange membrane fuel cell provides electric energy for airborne electric equipment of the engine; the problems that the electric power requirement of an existing turboshaft engine is gradually improved, power of storage batteries such as lithium batteries is insufficient, and electric power is difficult to extract from a power turbine are solved. Meanwhile, the ammonia working medium serves as power generation fuel, the energy conversion efficiency of the aviation turboshaft engine is improved, and energy consumption is reduced.

The turboshaft engine comprises a speed reducer, a gas compressor, a combustor and an exhaust nozzle, the speed reducer and the combustor are both fixedly connected with the gas compressor, the input end of the combustor is connected with the output end of the gas compressor, and the exhaust nozzle is fixedly connected with the output end of the combustor. The speed reducer comprises a case, a transmission assembly and an auxiliary assembly are arranged on the inner wall, the transmission assembly is connected with the combustor, and the auxiliary assembly is connected with the gas compressor. And the output end of the lubricating assembly is communicated with the interior of the case. The turboshaft engine has the core advantages that a transmission assembly and an auxiliary assembly are integrated in the same case, the overall size is greatly reduced, and the compact installation requirement of the turboshaft engine is met; the case provides reliable protection for internal parts, and interference of the external environment on the transmission precision is reduced. By optimizing a power transmission path, the transmission assembly specially receives power transmission of the combustor, the auxiliary assembly concentrates on driving the gas compressor and related auxiliary parts, power distribution precision is achieved, and the power transmission stability of the whole machine is remarkably improved.

The application discloses a turbo-shaft engine ground bench performance debugging method based on turbine guide vane area, comprising the following steps: obtaining different gas turbine guide vane areas A GT and power turbine guide vane areas A PT under the condition of same gas generator physical rotating speed t4.5 , corresponding gas temperature change data ΔT, power change data ΔP, and forming gas temperature change and power change tables respectively; obtaining different A GT and A PT under the condition of same take-off power g , corresponding specific fuel consumption change data Δsfc and rotating speed change data Δn, and forming specific fuel consumption change table and rotating speed change table respectively; obtaining A GT and A PT corresponding to the change data which simultaneously satisfy the related constraint conditions from the two-dimensional linear interpolation of each table, and taking A GT and A PT as candidate A GT and A PT ; and finding A GT and A PT with the minimum adjustment amount for performance debugging. The application can improve the debugging efficiency and reduce the debugging cost.

A turboshaft engine transmission with a torque measurement system is provided. The turboshaft engine transmission includes a front housing, a rear housing, a mating flange axially interposed between the front and rear housings, the mating flange defining a sensing bolt hole, assembly bolts disposed to assemble the front and rear housings to the mating flange with a preload sufficient to maintain assembly during engine operational conditions and a sensing bolt installed in the sensing bolt hole to measure forces exerted on the mating flange by the front and rear housings.

The present invention discloses a method for processing a corrosion-resistant blade adjustment ring. Two half-ring blanks arranged in half are obtained by machining, and after the two half-ring blanks are paired to process multiple end holes, a suitable first positioning bushing is selected by calculation, so that the first positioning bushing with clearance fit can effectively avoid dimensional deviation when the first ring body is decomposed, ensuring that the subsequent second positioning bushing with interference fit can still meet the technical requirements. After the mounting hole is machined, sulfuric acid anodizing treatment can be performed to effectively ensure the integrity of the anodized protective film on the surface of the blade adjustment ring, improve the corrosion resistance of the blade adjustment ring, and prevent the occurrence of corrosion to the greatest extent, so that the turboshaft engine using the blade adjustment ring of this scheme has high reliability and long service life. At the same time, the machining process after sulfuric acid anodizing treatment is reduced, and the risk of damaging the anodized protective film during subsequent machining is reduced. The method has strong practicality and is suitable for wide promotion and application.

A non-propulsive turbomachine (2) for an aircraft (M) fuel conditioning system (Q) comprising: a first compressor (21), a combustion chamber (23) supplied by a first airflow (A1) flowing in a general supply channel (30) connected to the first compressor (21), a second compressor (22) connected to the general supply channel (30) by a first sampling channel (31), so as to be supplied with a second airflow (A2) from the first compressor (21), and to a cabin air conditioning device (6), a single gas turbine (24) connected to both the first compressor (21) and the second compressor (22) by a single turbomachine shaft (25), a second sampling channel (33) connecting the general supply channel (30) and the cabin air conditioning device (6), and at least one sampling valve (4),mounted on the first sampling channel (31) and configured to allow or prevent the circulation of the second airflow (A2) in the sampling channel (31). Abstract figure: Figure 3,

The invention provides a modeling method and device for a turboshaft engine of a helicopter, and relates to the technical field of turboshaft engine simulation, and the method comprises the steps: sequentially solving the thermodynamic characteristics of all parts according to the position sequence of all parts in the turboshaft engine of the helicopter, and obtaining the modeling of the turboshaft engine. According to the method, the thermodynamic characteristics of all the components are solved in sequence according to the station sequence of all the components in the turboshaft engine, modeling of the turboshaft engine is achieved, it is guaranteed that the inlet parameter of each station is consistent with the outlet parameter of the previous station, and the problem that the matching performance between the parameters of all the stations is poor is solved.

The invention discloses a turboshaft engine thermodynamic cycle parameter evaluation method based on cost and task requirements, and the method achieves the evaluation of the thermodynamic cycle parameters of a turboshaft engine on the premise of comprehensively considering the coupling influence of cycle parameters, all parts and air system parameters. According to the method, the relation between the thermodynamic cycle parameters and all parts and the whole machine cost is established, and the distribution condition of the engine cost and the engine and fuel oil weight coefficient about the cycle parameters is obtained by combining the balance between the aircraft task and the weight, so that a reliable and scientific basis is provided for thermodynamic cycle parameter evaluation and further optimization; the method can be used for evaluating the applicability of different thermodynamic cycle parameter combinations to the flight platform, so that the cost of an engine is more advantageous, the minimum residual weight coefficient of the flight platform is still within an acceptable range, and the voyage of the flight platform is further improved.

The invention relates to an aircraft propulsion system (100), which comprises a turbine engine (102A, 102B) provided with a rotary portion (104A, 104B), a starter (122A, 122B), a freewheel clutch system (124A, 124B) and a control system (126) designed, when the turboshaft engine is operating at idling speed, to receive a standby request (Rv), and in response to the standby request (Rv), to turn off the turbine engine (102A, 102B). The control system (126) is designed, in response to the standby request (Rv), before the turbine engine is switched off, to control the starter so that the speed (ND) of the starter becomes non-zero, and after the turbine engine is switched off, to control the starter to maintain the non-zero speed (ND) of the starter during the decrease in the speed (NGG) of the rotary portion to the standby speed (NGGv).

This application discloses a method, apparatus, equipment, and medium for calibrating the torque of an aero-gas turbine shaft engine based on a torque meter. The method includes the following steps: S1, engine steady-state performance testing: starting the engine to idle, then slowly increasing the engine to several takeoff torque states, recording the torque values ​​corresponding to the power turbine shaft torque sensor and torque meter at each takeoff torque state; S2, torque calibration: selecting N=2 takeoff torque states for the engine, the torque values ​​corresponding to the power turbine shaft torque sensor and torque meter, and obtaining the calibration function relationship through linear fitting; S3, post-calibration check: calculating the calibrated torque values ​​of the power turbine shaft torque sensor at several takeoff torque states based on the calibration function relationship. If the differences between each calibrated torque value and the torque value corresponding to the torque meter all meet the requirements, then the torque calibration meets the requirements. This application improves the accuracy of engine torque calibration.

This application discloses a method, apparatus, device, and medium for dynamically optimizing turboshaft engine component characteristics. The method comprises the following steps: S1. Calculating typical engine parameters for each flight phase, iteratively obtaining multiple rounds of typical parameter calculation results by varying the relative conversion speed increment, and selecting the compressor characteristic diagram with the lowest weighted average fuel flow rate as the optimal optimized design result for the compressor; S2. Based on the optimal optimized design result for the compressor, performing engine parameter calculations to determine the positional relationship between the power turbine's highest efficiency zone on the turbine characteristic diagram and each aircraft phase, adjusting the power turbine efficiency at the design point using a double Gaussian correction formula, and iteratively obtaining multiple rounds of typical engine parameter calculation results for each flight phase. The power turbine characteristic diagram with the lowest weighted average fuel flow rate is selected as the optimal optimized design result for the power turbine. This application achieves low optimization cost and high efficiency and accuracy.

The invention relates to the technical field of turbofan engine fuel system control, in particular to a turbofan engine fuel pump current interference detection and extraction circuit and analysis device, the circuit comprises a voltage following module and a bidirectional threshold detection and extraction module; in the bidirectional threshold detection and extraction module, the positive input end of a first operational amplifier is grounded, the negative input end of the first operational amplifier is connected with the output end of a voltage following module, and the output end of the first operational amplifier is connected with the cathode of a first diode and the anode of a second diode. The anode of the first diode is respectively connected with the output end of the voltage following module and the negative input end of the second operational amplifier; the cathode of the second diode is respectively connected with the output end of the voltage following module and the positive input end of the second operational amplifier; and the output end of the second operational amplifier is connected with an external current interference analysis module. Therefore, through the hardware circuit, the detection and extraction effect of the interference current can be effectively improved.

This invention belongs to the field of aero-engine turboshaft technology and discloses an automatic monitoring method and system for the performance degradation of aero-engine turboshaft engines. The method includes: determining the number of monitoring states and the altitude range corresponding to each monitoring state; determining the converted power value for each monitoring state; determining the turboshaft engine power; determining the converted gas generator speed and converted turbine inlet temperature corresponding to the converted power value for each monitoring state; determining the variation trend of the converted gas generator speed and converted turbine inlet temperature with the number of sorties for each monitoring state, and characterizing the degradation trend of the turboshaft engine through these trends. This invention does not require baseline performance, can monitor the engine performance variation trend with the number of sorties, and can monitor the performance degradation of the engine over operating time, without adding extra maintenance workload.

HIGH-STABILITY, HIGH-AGILITY THREE-SAILING SURFACE AIRCRAFT The three-sail aircraft (1), (2), and (3) is equipped with movable flaps (6), (7), and (8) on each of the lifting surfaces. The flaps of the upstream lifting surface (6) are connected to the flaps of the main wing (7) by a rigid linkage with a deflection angle ratio of 50%. The forces on the downstream lifting surface (horizontal tail) (3) reverse between takeoff (negative lift) and cruise (positive lift). Stabilizing fins (5) are positioned at the rear of the fuselage in the lower section at an angle of 30° to the aircraft's centerline. A weighted control rod (22) located in the vertical tail (4) provides stabilization for the aircraft in the presence of severe turbulence.The propulsion system is located at the rear of the fuselage and consists of either a piston engine driving a multi-bladed propeller, one or two turboshaft engines driving a multi-bladed propeller or two contra-rotating propellers, a turbojet engine, or one or two electric motors driving a multi-bladed propeller or two contra-rotating propellers. This aircraft can be piloted either by a human pilot on board or by an electronic computer controlled from the ground or autonomously.