Single Engine Failure Research Articles

Research on attitude control strategy of single engine failure in the first-stage flight phase of new generation launch vehicle

The successful launching of the new generation of launch vehicle is of great practical significance to the development of China Space, and the cryogenic liquid engine commonly used in the launch vehicle has a certain probability of failures, which may affect the success or failure of rocket launching. In this paper, a Bolza problem optimization algorithm based on Radau pseudo-spectral method is proposed to solve the problem of single engine failure in the first-stage flight phase of the new generation launch vehicle by simulating fault injection and taking the attitude angle as optimal control quantity. The simulation results show that this method can effectively eliminate the influence of single engine failure on orbit injection accuracy, achieve the fault absorption of new generation launch vehicle engine, improve the fault tolerance of attitude control system, and further guarantee the system reliability.

Whirl flutter optimisation-based solution of twin turboprop aircraft using a full-span model

Whirl flutter is a specific type of flutter instability, relevant for turboprop aircraft, caused by the effect of rotating parts as a propeller or a gas-turbine engine rotor. The proposed optimisation-based analytical procedure is used to determine the critical values of the engine attachment stiffness parameters for the preselected flutter speed. For the half-span model, two design variables are used. The objective function is defined as the minimization of the engine vibration mode frequency sum. Design constraints keep the engine frequency ratio and the flutter stability at the selected velocity. However, application of a full-span model is necessary in some cases. In this case, special models of both symmetric and antisymmetric engine vibrations and four design variables must be used. Design constraints maintain the pitch mode frequency ratio, the yaw mode frequency ratio and the critical mode frequency ratio. Critical modes are dependent on the relation between the rotational direction of both propellers (identical or inverse). A flutter design constraint is applied as well. The described methodology is demonstrated on the application example of a twin-engine commuter aircraft. Demonstrated cases include symmetrical revolutions of propellers for both identical and inverse directions of rotation, cases of single engine failure and single propeller feathering, and finally, cases of unsymmetrical revolutions including the reduced and increased revolutions of a single propeller, for both identical and inverse directions of rotation.

Abort recovery strategy for future vertical landing systems

This paper summarizes the abort recovery strategy for future vertical landing spacecraft in case of single engine failure, by applying state-of-art non-linear control law.The abort recovery strategy applied to future vertical landing systems with multiple engines are composed of two approaches. First is to shutdown the counter engine of the failed engine in order to minimize the attitude perturbation. And the second, which is the key approach, is applying a nonlinear control law. Such an approach has the robustness to the large attitude motion, and also to the variation in weight, moment of inertia and center of gravity. The control law applied has been developed under the collaboration with Nagoya University. To demonstrate the above mentioned strategy, Mitsubishi Heavy Industries is preparing flight test by a subscale demonstrator called “OEEX”, which stands for “One Engine inoperative EXperimental vehicle”. The results of application of such strategy are described. Also, the status of the OEEX is presented.

수직이착륙기의 착륙접근시 단일엔진고장 및 비행전이 영역 해석

본 연구에서는 착륙 접근시 단일엔진 고장 분석 및 비행전이 해석을 통하여 수직이착륙기에서 요구하는 최적의 엔진추력과 날개하중비를 구하고자 한다. 항공기 해석 모듈들은 기존의 항공기 통합 사이징 프로그램을 기반으로 하고 있으며, 항공기 설계 및 해석을 하나로 묶는 컴퓨팅 체계 구축은 사용자 용도에 맞게 주문 제작이 용이한 반 완성 프로그램인 EMDIOS를 설계프레임웍으로 사용하였다. 시뮬레이션 결과, 고도 40 ft 근처가 단일엔진고장시 가장 위험한 영역임을 확인할 수 있었으며, 비행전이 영역에 대한 파라메타 연구를 통하여 엔진 추력은 클수록, 반면에 엔진틸트각 회전속도와 날개 하중비는 작을수록 비행전이시의 고도 상실이 작아짐을 알 수 있었다. The objective of this study is to find the optimal thrust condition and wing loading of a vertical take-off and landing (VTOL) fixed-wing aircraft through a single engine failure analysis during landing approach and an analysis of transition flight. The aircraft analysis modules used in the study are based on the aircraft synthesis program. To achieve the computing infrastructure for aircraft design and analysis, the EMDIOS was employed as a design framework, which is a semi-completed application program and ready to customize. Simulation results reveal the most critical height at the event of single engine failure is approximately 40 ft. And, in order to avoid a significant loss in altitude during the transition, the thrust to weight ratio must be kept high, while both the engine tilt speed and the wing loading must be kept low, as confirmed by the analysis results.

Optimal City-Center Takeoff Operation of Tiltrotor Aircraft in One Engine Failure

This paper investigates optimal tiltrotor flight trajectories and performance in vertical takeoff operations from a city-center heliport considering the possibility of one engine failure. A two-dimensional longitudinal rigid body model of a tiltrotor aircraft is used. Tiltrotor flights after engine failure in both continued takeoff and rejected takeoff are formulated as nonlinear optimal control problems that minimize the heliport size, subject to various constraints from safety considerations and tiltrotor performance limitations. These problems are parametrized via collocation into parameter optimization for numerical solutions. Extensive numerical solutions are obtained, and sensitivity analyses are conducted to examine effects of model parameter uncertainties. Optimization results indicate that the maximum gross weight capability of a tiltrotor in a vertical operation is determined by the need for safe landing in the event of a single engine failure. The height at which the decision for either continuing or rejecting the takeoff is made depends on the need to maintain sufficient height during the continued takeoff flight after an engine failure.

Optimal Short Takeoff of Tiltrotor Aircraft in One Engine Failure

Optimal tiltrotor flight paths in the event of a single engine failure during short takeoff operations are investigated. A vertical plane rigid-body model is used that has as state variables aircraft position, body components of aircraft velocity, pitch angle and rate, and rotor angular speed. The control variables are the thrust coefficient of one rotor, the pilot's longitudinal stick displacement, and the nacelle angle. The model uses both parameters and aerodynamic data of the XV-15 research aircraft. The tabular aerodynamic data are interpolated with smooth functions. Tiltrotor flights after engine failure are formulated as nonlinear optimal control problems. Both continued takeoff and rejected takeoff flight after an engine failure are studied. Performance indices are selected to minimize runway length, subject to various constraints from safety considerations and tiltrotor performance limitations. These optimal control problems are parameterized via collocation into parameter optimizations for numerical solutions. Extensive numerical solutions are obtained, and sensitivity analyses are conducted to examine effects of model uncertainties

Optimization of Balanced Field Length Performance of Multi-Engine Helicopters

The development is described of a balanced field length trajectory optimization approach for a multi-engine helicopter sustaining a single engine failure. Whereas most studies typically focus on a single flight phase, this study is based on a multiphase formulation, in which the rejected takeoff and continued takeoff trajectories are optimized simultaneously, subject to a field length balancing constraint. The advantage of this approach is that, for any given engine failure time, it allows the flight phase where all engines are still operating to be optimized in such a way that the solution represents the best possible compromise between the conflicting requirements set by the rejected takeoff and continued takeoff flight phases. In addition to balanced field length calculations, the optimization of unbalanced rejected takeoff has been addressed. Combined considerations of balanced and unbalanced rejected takeoff give insight in the choice of the critical decision point. The most important result of the overall optimization process is the optimal all-engines-operating takeoff flight path up to the critical decision point

Optimal vertical takeoff and landing helicopter operation in one engine failure

This article presents optimal vertical takeoff and landing (VTOL) trajectories of a multiengine helicopter in the event of a single engine failure. A point-mass model of the UH-60A helicopter is used. Time derivatives of thrust coefficients are treated as control variables. A first-order response dynamics is assumed for the contingency power available after an engine failure. Ground effect is included in the study of landing trajectories. A linear backup procedure is assumed for normal takeoff and a straight-in procedure for normal landing. Nonlinear optimal control problems are formulated for flights of rejected takeoff, continued takeoff, continued landing, and balked landing after one engine failure. Constraints on rotor speed and thrust vector are included. In rejected takeoff and continued landing, the cost functional minimizes the distance between the touchdown point and the original takeoff point, subject to impact speed limits at touchdown. In continued takeoff and balked landing, two cost functionals are used subject to specified final speed components required for a steady climb. One minimizes the horizontal distance required to achieve the steady climb, whereas the other maximizes the minimum altitude during the transitional flight from engine failure to the steady climb. Extensive numerical solutions are obtained. In all optimal trajectories, the helicopter adjusts its power requirement to accommodate the contingency power available. The helicopter gross weight capability hi VTOL operation is limited by safe rejected takeoff or continued landing. A concept of balanced-weight takeoff and landing decision point is proposed from which the maximum gross weight hi emergency landing and in flying-out maneuver is the same.

Critical considerations for helicopters during runway takeoffs

This article presents optimal runway takeoff trajectories of a multiengine helicopter in the event of one engine failure. A point-mass model representative of the UH-60A helicopter is employed. Pilot response delays are ignored. Two optimal control problems are formulated for both continued takeoff and rejected takeoff after a single engine failure. Subject to specified safety conditions, the first formulation minimizes the runway length required for a given takeoff weight, whereas the second maximizes the takeoff weight for a given runway length in continued takeoff and for an unspecified runway length in rejected takeoff. Constraints on thrust angle, and thrust magnitude are included. For the continued takeoff, an optimal choice of the takeoff safety speed is determined to be about 15 ft/s above the initial airspeed at engine failure. Maximum takeoff weight is dictated by the remaining power after one engine failure if the runway length is sufficient, and is determined by the available runway length otherwise. For the rejected takeoff, the minimum runway length is a function of initial conditions at engine failure, and is roughly independent of takeoff weights. These results replicate some key characteristics of those from flight tests reported in the literature.

Development of the iterative guidance mode with its application to various vehicles and missions.

The guidance errors at orbital insertion or lunar conic injection in the position magnitude, velocity magnitude, and the path angle, with respect to the specified terminal conditions that result from three-sigma vehicle perturbations including a single engine failure in the S-IC or S-II stages are: AR = ±20 m, AF = ±0.005 m/sec, and AT = ±0.001°. The velocity error at cutoff is due to the inability of the vehicle to achieve cutoff at the proper millisecond. The errors at lunar conic injection, assuming perfect navigation, result in less than 100 km error in attaining the specified radius of closest approach to the moon if no midcourse corrections are made. The basic criterion for evaluating the performance of the Saturn V navigation system is the midcourse correction velocity required to achieve the desired radius of lunar approach at the desired time. Table 1 summarizes the corrections required at two hr after injection onto the lunar trajectory. The data are based on a stay-time of 2.7 orbits. Environmental and propulsive uncertainties are included with the equation errors. The midcourse velocity requirements are well within the available velocity budget for a Saturn-V Lunar Landing Mission; the major portion of the budget is used to correct the platform hardware errors.