Engine Thrust Research Articles

Research on operating characteristics of single-spool turbojet enginewith the afterburner AL-21F

The single-spool, single-thread, turbojet engine with afterburner AL-21F has been used on fighter aircraft for a long time, but there are few studies on this engine in Vietnam. This article uses the theoretical approach, calibrates according to the technical description and actual operation, combined with specialised software such as GasTurb and GSP to calculate engine thermodynamic parameters in some actual flight conditions. The obtained results showed that, when the flight speed increases, the specific fuel consumption increases linearly, while the initial engine thrust decreases and reaches the minimum value at M=0.3÷0.5, then continue to increase and reach the maximum at the speed limit. When increasing the height, the speed and regulating spool speed of the engine remain the same, the engine thrust decreases as well as the specific fuel consumption increases. When increasing the rotor rotation speed, the engine thrust increases linearly but the specific fuel consumption decreases gradually and remains almost the same at engine rpm above 85%. The results allowed for a better understanding of the operating process of the AL-21F engine in particular and turbojet engines in general or used to build simulation models operating close to reality.

Characterization of pintle engine performance for GO2/kerosene propellants

Liquid rocket engine with a wide range of thrust variations has drawn much attention in recent years. A GO2/kerosene deep-throttling variable thrust engine using a liquid–gas pintle injector is presented and tested with aim of characterizing engine performance, regulating capability, and heat flux distribution. Three-dimensional numerical simulations are conducted to investigate the flame appearance and spray characteristics of the engine. The SST k-omega model is used for modeling turbulence, the Wave model is used for modeling the secondary breakup, and a multi-step quasi-global combustion mechanism and eddy dissipation concept model are used for modeling combustion. The GO2/kerosene pintle engine achieves stable combustion at 0.18 ∼ 4.35 MPa with a thrust of 21.30 ∼ 864.70 N. The thrust ratio of the pintle engine is 40.6:1 indicating a deep-throttling variable thrust capability. The thrust regulation of the pintle engine using the closed-loop control has high accuracy and response speed, as well as high disturbance suppression capability. The effects of combustor pressure, spray characteristics, and reactants concentration distribution on flame appearance and heat flux distribution of the engine are studied in detail. The engine equipped with a liquid–gas pintle injector could ensure a good atomization effect and high combustion efficiency by actively adjusting the propellant injection pressure drop coefficient. This work will provide a reference for the design of the pintle injector and thermal protection system of deep-throttling variable thrust rocket engines.

QCr0.8 Cu alloy /S06 stainless steel bimetal structure via In718 multi-interlayer fabricated by laser powder hybrid additive manufacturing

Fabricating high strength high conductivity (HSHC) Cu alloy inner parts by laser powder bed fusion (LPBF) and high-strength steel out layers on Cu alloy surfaces by laser-direct energy deposition (LDED), e.g., laser powder hybrid additive manufacturing (LPH-AM), has shown promise in the development of rocket engine thrust chambers with large thrust forces. In this study, we successfully fabricated a QCr0.8 HSHC Cu alloy/S06 stainless steel bimetal structure via In718 multi-interlayer using LPH-AM. We investigated the relative density, defects, microstructure, and interfacial characteristics of the bimetal structure in detail. The results indicate that the sample achieved up to 99.9% relative density with no naked defects and good metallurgical bonding at the QCr0.8/In718 and In718/S06 interfaces. The LPH-AM-ed QCr0.8/In718 interface exhibited a combination of columnar and equiaxed dendrites. The vertically combined sample exhibited an average ultimate tensile strength (UTS) and break elongation (EL) of 300.3 ± 10.6 MPa and 15.0 ± 1.4%, respectively, with fracture occurring at the QCr0.8 side. This study provides a new method for fabricating complex-shaped Cu/steel bimetal components, such as rocket engine thrust chambers.

Research on afterburning control of more electric engine with a nonlinear fuel supply system

More electric engine (MEE) with electric pump metering fuel has the problem of unstable engine control due to the dead-zone characteristics of afterburner fuel actuator. On the basis of conventional μ modified adaptive control method, a compound modified adaptive control (CMAC) system based on neural network inverse model is proposed in this study. At first, the afterburner fuel actuator model of MEE is established through Simulink/AMESim co-simulation to study its fuel supply characteristics and reveal the reasons for dead-zone characteristics of the actuator. Subsequently, an integrated model of afterburner fuel actuator/MEE is established to investigate the effect of dead-zone characteristics on engine performance. Finally, the CMAC system for afterburner fuel flow closed-loop control based on neural network inverse model and improved μ modified adaptive control method is designed. The simulation results show that the dead-zone characteristics of afterburner fuel actuator can cause engine shaking and unstable operation. The engine thrust has a continuous fluctuation with amplitude of about 0.2%. The proposed CMAC system can effectively avoid the dead-zone interval of actuator and achieve stable transition of the engine state in the entire operating range of the afterburner fuel actuator. Thus, the stability of the control system can be significantly improved. Meanwhile, the improved μ modified adaptive control algorithm in the CMAC system has better dynamic characteristics when compared with the conventional μ modified adaptive control algorithm.

Improved Thrust Performance Optimization Method for UAVs Based on the Adaptive Margin Control Approach

This study proposes a strategy for improving the thrust performance of fixed-wing UAV turbine engines from the perspective of aircraft/engine integration. In the UAV engine control process, the inlet distortion caused by the angle of attack change is taken into account, the inlet distortion index is calculated in real time by predicting the angle of attack, and the influence of the inlet distortion on the engine model is analyzed mechanically. Then, the pressure ratio command is adjusted according to the new compressor surge margin requirement caused by the inlet distortion to finally improve the engine thrust performance. To verify the effectiveness of the algorithm, an adaptive disturbance rejection controller is designed for the flight control of a fixed-wing UAV to complete the simulation of horizontal acceleration. The simulation results show that, with this strategy, the UAV turbofan engine can improve the turbofan engine thrust performance by more than 8% under the safety conditions.

Numerical Investigation on the Service Life of a Liquid Rocket Engine Thrust Chamber

Rocket engine thrust chambers withstand very high temperatures and thermal gradients during service that induce multiple damaging phenomena such as plasticity, low-cycle-fatigue (LCF) and creep. Numerical models can be used during the design of these mechanical components in order to simulate the main mechanical damaging processes, accounting for complex material behavior as due to non-linear hardening phenomena and viscoplasticity. This work represents an improvement upon previous research by the authors, with particular reference to the addition of the Wang–Brown fatigue criterion, to consider the effects of multiaxiality and non-proportionality of loads, and the Voce model to account for non-linear isotropic hardening. A precipitation hardened copper alloy has been considered as the material of the thrust chamber internal structure. The most critical areas resulted to be on the internal surface of the chamber and in particular in correspondence of the throat region, consistently with experimental tests available in the literature conducted on similar geometries. Results demonstrated that low cycle fatigue and thermal ratcheting (plastic instability) are the dominant damaging phenomena for the considered test case.

Performance assessment for a novel supersonic turbine engine with variable geometry and fuel precooled: From feasibility, exergy, thermoeconomic perspectives

Aiming at the problem of turbine engine insufficient thrust caused by low efficiency at high Mach number, a three-axis variable geometry dual-mode turbine engine scheme based on ammonia and n-decane equivalent dual fuel precooling and adjustable low-pressure (LP) turbine guide vane is proposed, and parameters such as specific thrust price, relative thrust economic ratio, volumetric specific fuel consumption are defined to scientifically evaluate the feasibility of the scheme. The ammonia and n-decane dual fuel method is used to expand the efficient working range of the precooler. On the one hand, in the low speed section, the engine operates in the turbofan mode, adjusting the LP turbine guide vane can offset the adverse impact of the precooler low efficiency on the engine thrust performance under the low inlet total temperature, and improve the engine economy. On the other hand, the high speed section operates in the turbojet mode, and the equivalent fuel precooling combined with the LP turbine guide vane adjustment is conducive to improving the working efficiency and thrust performance of the turbojet engine. Feasibility, exergy and thermoeconomic analysis indicate that under different operating modes, due to different control plans, the influence of fuel ratio and heat transfer efficiency changes of dual fuel precooler on engine performance is opposite. For the typical climbing process with H = 0, Ma = 0 to H = 30 km, Ma = 5, this scheme can significantly increase the engine thrust, with a maximum increase of about 41%, but it increases the volume fuel consumption, reduces the engine exergy efficiency, exergy sustainability, and thermoeconomic under the turbofan mode. Therefore, the variable geometry turbine engine scheme based on dual fuel precooling and adjustable LP turbine guide vane has high application potential in supersonic commercial flight under the premise of slightly losing economy.

Flutter and divergence instability regions of a column subjected to the load realized by means of a jet engine with an additional guiding structure

In this paper, a new slender system is considered — a column that is subjected to a compressive follower load. The studied structure uses a system of rigid transoms guiding the loaded end of the column with the limitation of their mutual rotation by means of rotational springs. This additional element does not transfer the compressive load but affects guidance of the loaded end of the column. In the presented form, the solution used to guide the end of the column was not previously taken into account. The boundary problem was formulated on the basis of Hamilton’s principle, taking into account (due to the non-linearity) the small parameter method. The theoretical, numerical, and experimental studies were conducted by determining an influence of the system parameters (in particular the element guiding the loaded end) on the instability type. On the basis of the kinetic stability criterion, critical divergence and flutter loads as well as limit values of parameters in the system were determined, at which instability occurs. The research was limited to the determination of the first linear component of vibration frequency. As a part of numerical research, detailed considerations were made on characteristic curves. It has been shown that with an appropriate selection of the parameters of the guiding system, a new shape of the characteristic curves can be obtained. This new feature shows that in the boundary case (instability change from flutter to divergence) the critical divergence load is greater than the critical flutter. In order to confirm the correctness of the mathematical model and thus the results of numerical tests concerning the occurrence of flutter and divergence areas the experimental tests were carried out consisting of recording the behavior of the system with a quick camera when it was loaded with a model jet engine. The correctness of the shape of characteristic curves was confirmed by the experimental frequency analysis of considered columns at different jet engine thrust magnitudes. The conducted research may be of great importance in the control of the flutter phenomenon of slender systems exposed to follower forces, including those resulting from the action of the wind.

Diagnostics using a physics-based engine model in aero gas turbine engine verification tests

In the present study, the application of a diagnostic method using a physics-based model is proposed during engine verification testing. A physics-based engine model is constructed, based on a 0-dimensional system analysis method, and then revised through performance adaptation with a basic engine model. Adaptation factors are used for the compressor, secondary air system, nozzle, exhaust gas temperature, and engine thrust. Gas path analysis is conducted using the physics-based engine model and the performance test data in real time. A health parameter that can quantitatively evaluate the performance difference of components was defined. Health parameters for the compressor capacity and efficiency, secondary air system, and nozzle are calculated from the gas path analysis. The multi-variable Newton–Raphson method is employed for the performance adaptation and gas path analysis in real time. A single-shaft gas turbine engine, which is constructed based on the core part of a turbofan engine currently under development, is utilized to test and evaluate the effectiveness of the proposed method. Three tests are performed for engine verification. The proposed diagnostic method is applied to the aero gas turbine engine verification test. In 33 test cases, the fault detection rates for the compressor's variable guide vane, secondary air system, and variable area nozzle are 100%, 91%, and 79%, respectively. As a result, it is confirmed that the proposed method can simultaneously detect abnormalities in the compressor's variable guide vane, secondary air system, and variable area nozzle. In addition, this study indicates that applying the physics-based engine model to the diagnostics logic can effectively detect failures with a small amount of data.

A Nonlinear Programming-Based Morphing Strategy for a Variable-Sweep Morphing Aircraft Aiming at Optimizing the Cruising Efficiency

This work develops a morphing decision strategy to optimize the cruising efficiency for a variable-sweep morphing aircraft, and a simple and practical guidance and control system is given as well. They can work in tandem to accomplish a cruise mission effectively. To make the morphing decision accurately, we take into account the equilibrium equations of forces; the variations in airspeed, altitude, and mass and the optimal configurations for different cruise conditions are solved based on the nonlinear programming method with the objective of minimum engine thrust. Considering that a large amount of computational resources are required to solve the nonlinear programming problems, we establish an offline database of the optimal configurations and design a database-based online morphing decision process. In addition, the proposed morphing decision strategy includes an anti-disturbance mechanism, which ensures that the optimal configuration can be given accurately without chattering under fluctuating airspeed measurements. Comparative results from the simulations finally validate the effectiveness of the proposed strategy.

Research of Comprehensive Modeling of Two-Dimensional Vector Nozzle/Actuator Load/Turbofan Engine

To investigate the effect of the dynamic deflection process of a two-dimensional vector nozzle on its actuator load and engine performance, a comprehensive modeling method of two-dimensional vector nozzle/actuator load/turbofan engine is proposed. Firstly, based on the two-dimensional vector nozzle spatial motion model, combined with force and moment analysis, the mechanical model of vector nozzle actuation system is established; The load spectrum of the actuating system under different working conditions and deflection angles is obtained by CFD numerical simulation results; Using AMESim, the hydraulic actuation system model of two-dimensional vector nozzle is established. Finally, a two-dimensional vector nozzle/actuator/turbofan engine comprehensive model is established based on the component-level model of turbofan engine. The comprehensive model can accurately calculate the changes of main performance parameters such as engine rotor speed, thrust and pre-turbine temperature when the nozzle vector deflects. The simulation results show that the load on the nozzle actuator is positively correlated with the deflection angle, and the load change will lead to the fluctuation of the output displacement of the hydraulic actuation system. During vector deflection, due to the change of engine circulation capacity and thrust coefficient, the performance parameters such as engine low-pressure rotor speed and surge margin fluctuate greatly.

Identification method of temperature corrections to aircraft engine thrust in flight tests

The article is devoted to identifying the dependence of thrust on the outside air temperature for a turbojet engine as part of an aircraft. A methodic that allows to determine temperature corrections to the effective thrust of a turbojet engine at fixed operating modes during flight tests is presented. The application of the developed methodology in flight testing of a high-speed high-altitude aircraft is shown. A distinctive feature of the method is the calculation of the engine thrust increment through the excess thrust coefficient at zero lift in straight-line horizontal acceleration modes. Keywords aircraft, aircraft testing, turbojet engine, thrust identification, turbojet engine characteristics

Способ оценки адекватности результатов численного моделирования поля параметров неравномерного потока на входе в вентилятор ТРДД экспериментальным данным

The article substantiates the need to intensify the process of oil cooling in the oil systems of aircraft turbojet engines. Currently, an increase in the thrust of turbojet engines is achieved by increasing the air pressure ratio in the compressor and the gastemperature in front of the turbine. However, an increase in thevalues of theseparameters leads to an increase inthe thermal loadof theengine elements,including rotor supports, which causes an increase in oil temperature. The oil systems of modern turbojet engines still use the same oils as several decades ago.Therefore, the intensification of the oil cooling process, which allows reducing its temperature, is the main way to preserve the lubricating properties.It is proposed to intensify the oilcooling process in the oil system of the bypass turbojet engine byinstalling an air-oil heat exchanger in the bypass duct.A thermal calculation of fuel-oil and air-oil heat exchangers has been carried out, the results of which show large values of the heat transfer coefficient when the oil is cooled by the bypass duct air

A simplified one-dimensional transient heat transfer model for rocket thrust chamber

The heat transfer model is essential to rocket engine thrust chamber, for it determines the accuracy of chamber wall temperature and heat flux distribution. In order to predict the transient temperature field during startup and shutdown, we introduce a simplified one-dimensional model of transient heat transfer for thrust chamber. The heat transfer process is divided into two parts: the convective heat transfer from hot gas to the inner wall and the heat conductance through the chamber wall. The Bartz formula modified with constant coefficient C is used to characterize the convective heat transfer and the one-dimensional transient heat transfer model of slab is the calculative foundation of the heat conductance. Finally, we propose a strategy to iteratively solve the model. To verify the model, it is implemented in a series of hot run experiments for a LOX/Methane heat sink thrust chamber. The average error of wall temperature evaluated at different locations of the thrust chamber inner wall can be reduced up to a 1.76%. The application of our novel method provides an accurate and efficient tool to calculate chamber wall temperature and heat flux distribution in the thrust chamber design. In addition, through inverse analysis of the proposed model, we can also estimate the near-wall gas temperature, which is difficult to obtain in a hot run test, by acquiring wall temperature data.

Расчетное исследование продольной устойчивости квантомобиля

The purpose of the study is to elaborate a methodological evaluation of the car's rollover moment caused by the quantum engine thrust, in order to preserve the longitudinal stability of the quantomobile. In view of the lack of any full-scale data regarding the quantomobile concept under consideration, it appears to be relevant to obtain estimates of the significance of the stated problem, and this is the purpose of this paper. In the Matlab environment, there was constructed a model of normal reactions` distribution on the quantomobile supporting wheels. There was formed a research methodology based on the considered model. Based on the model of the quantomobile, an analogue of a truck, there has been conducted a computational study of the influence of the location of the QuEn on the longitudinal stability of the vehicle. The scheme of the forces operating on a quantomobile should become generic one for the formation of schemes of quantomobiles with additional thrusters for vehicle longitudinal stability increase.

Control Programs of the Aviation Gas Turbine Engine in the Modes of Acceleration, Gas Reset, Start-up. Optimization and Estimation of the Quality of Control Programs

The article is devoted to the formation of requirements for the accuracy of regulation of an aviation gas turbine engine, one of which is the maintenance of engine thrust at a given operating mode, regardless of the engine condition, within the gas temperature margin. Its value should not be significantly affected by turning on or off additional power and air consumers, as well as various regulatory influences on the part of the automatic control system (turning on or turning off the bypass in the compressor and blowing the housings, partial restriction of the supply of cooled air, changing the position of the guiding devices). Fulfilling the requirements for the accuracy of regulation is important for ensuring the reliability and safety of the operation of the power plant and the convenience of controlling the aircraft. In order to reduce operating costs, it is necessary that during operation, a minimum number of additional settings of the ACS in the acceleration mode, gas reset and start-up mode are required. The control program is implemented in the form of a automatic control system (ACS), which is a closed circuit of the main feedback. There is also a flexible local feedback loop in the circuit, which is designed to stabilize the ACS, which contributes to the fact that the ACS is quite stable. The presence of feedback in the ACS indicates that the system may be unstable, so the analysis of the ACS should include an assessment of its stability and, if necessary, the selection of measures and means for its stabilization. Changing the input signal at the first moment of time leads to a corresponding increase in deviation, since the links in front of the object and the object itself have inertia and therefore the rotation frequency cannot change instantly. The change in deviation, being an amplified amplifier, thyristor converter and generator, taking into account their inertia, leads to a gradual change in the control value, the voltage on the anchor, which smoothly changes the frequency of rotation of the shaft so that the tracking error, that is, the deviation, is directed to zero. Voltage feedback stabilizes the ACS and increases its speed. This is how tracking is done. The implementation of wireless technologies in the ACS of the gas station has been studied, which will allow to reduce the mass and dimensions of the nodes due to the reduction of the number of connectors and cables, increase the reliability and accuracy of the adjustment of the ACS, reduce maintenance costs and increase fire safety.

СИНЕРГЕТИЧЕСКИЙ СИНТЕЗ АВТОПИЛОТА ДЛЯ УПРАВЛЕНИЯ ДВИЖЕНИЕМ САМОЛЕТА-АМФИБИИ БЕ-200 ВО ВРЕМЯ ЗАБОРА ВОДЫ

The article is devoted to solving the problem of controlling the Be-200 amphibious aircraft whentaking water in the gliding mode. The specified mode of flight of an amphibious aircraft (AA) is one ofthe most intense for piloting. Controlling an aircraft in gliding mode requires the highest concentrationfrom the pilot, since the hydrodynamic characteristics decrease, but the aerodynamic ones increase,which causes situations in which the SA becomes uncontrollable. To ensure the safe operation of the AAduring water intake, as well as to ensure a steady takeoff, the pilot is forced to constantly maintain acertain value of speed, trim angle, and engine thrust by moving the engine control stick (ECS). In thisregard, it seems very appropriate to use the technical means of the autopilot, which provides control ofthe speed, engine thrust and trim angle to ensure the stability of the movement of the SA during the intakeof water. The article deals with the problem of synthesizing an autopilot to control the movement ofan amphibian aircraft during water intake, and at the same time the following tasks are solved: thesynthesis of an autopilot to control the longitudinal movement of an amphibian aircraft, as well as controllanding on water, followed by water intake and takeoff from the water. The substantiation of thechoice of a mathematical model of the movement of an aircraft on water is given. A procedure for thesynergistic synthesis of the autopilot controller of an amphibious aircraft in the water intake mode presented.The novelty of the ongoing research lies not only in considering the stable movement of an amphibiousaircraft during its operation on the water surface, but also in modeling the process of water intake in planning mode, which is a more complex task from the point of view of synthesis. As a result,the control laws for the autopilot obtained, which ensure the stable movement of the amphibious aircraftwhen descending along the glide path, landing on water, taking in water and subsequent takeoff withdamping of longitudinal oscillations along the trim angle throughout the entire water intake cycle.

Engine thrust model determination and analysis using a large operational flight database

Different engine thrust models are developed from operational flight data with limited a priori knowledge as part of a novel process for aircraft flight performance model determination. The given big data problem is solved by application of fundamental engineering knowledge and a specific data evaluation strategy. The resulting smart data approach is fundamentally different from existing artificial intelligence methods to solve such big data problems. A linear, a local-linear and a complex nonlinear thrust model are determined on the example of a given large database of operational flights with Airbus A 320neo aircraft. Even with limited information about the actual engine thrust from the available data, the resulting models allow to (well) predict the engine thrust characteristics within the required flight envelope. In addition, a temperature correction is predicted for the thrust model results to further enhance the model’s accuracy. Finally, the characteristics of the different thrust model implementations, evaluation results and thrust prediction quality are discussed.

Optimal aircraft arrival scheduling with continuous descent operations in busy terminal maneuvering areas

Continuous Descent Operation (CDO), an aircraft arrival procedure in which an aircraft descends from an optimal position with minimum engine thrust, can significantly reduce aircraft noise, fuel burn and emissions. The implementation of CDOs, however, requires more separation buffers among flights than traditional stepwise descents, resulting in loss of airspace and runway capacity. It is therefore challenging to maximize the number of aircraft performing CDOs for an airport, especially in busy terminal maneuvering areas. In this paper, we tackle the difficulty by developing an aircraft arrival scheduling model to sequence arrival aircraft before their merge point and generate conflict-free trajectories by means of a lateral path stretching method. The contributions of our study to the research community are twofold: firstly, we investigate a new aircraft arrival scheduling problem that incorporates both the CDO trajectory optimization and aircraft scheduling, and develop the first model simultaneously considering these two procedures; secondly, we conduct extensive computational experiments based on historical data in Guangzhou Baiyun International Airport, which would provide an important reference for future research in the area. Our numerical results demonstrate the benefit of the proposed CDO-based aircraft arrival scheduling approach in terms of reducing total flight time and saving fuel consumption when applied in this congested airport.

Life Analysis of Reusable Liquid Rocket Engine Thrust Chamber

The thrust chamber’s inner wall suffers high temperature and pressure differences from the coolant channel, which limits the life of the rocket engine. Life prediction of the thrust chamber really plays an important role in reusable launch vehicle propulsion systems. The Porowski beam model is widely used in the life prediction of reusable liquid rocket engine thrust chambers, which calculates the life caused by fatigue, creep, and thinning after each firing cycle. In order to analyze the life of the thrust chamber, a LOX/Kerosene rocket engine is investigated in this paper. The life analysis consists of pressure and temperature differences and structural parameters. Two kinds of inner wall materials were chosen for comparison in this research: OFHC copper and Narloy-Z alloy. The results are presented to offer a reference for the design and manufacture of reusable rocket engine thrust chambers in the future.