Thrust Drop Research Articles

Damped Iterative Explicit Guidance for Multistage Rockets with Thrust Drop Faults

A damped iterative explicit guidance (DIEG) algorithm is proposed to address the problem of the insufficient convergence of classical explicit guidance methods in the event of thrust drop faults in multistage rockets. Based on the iterative guidance mode (IGM) and powered explicit guidance (PEG), this method is enhanced in three aspects: (1) an accurate transversality condition is derived and applied in the dimension-reduction framework instead of using a simplified assumption; (2) the Gauss–Legendre quadrature formula (GLQF) is adopted to increase the accuracy of the method by addressing the issue of excessive errors in calculating thrust integration using linearization methods based on a small quantity assumption under fault conditions; and (3) a damping factor for solving the time-to-go is introduced to avoid the chattering phenomenon and enhance convergence. A numerical simulation was conducted in single- and multistage mission scenarios by gradually reducing the engine thrust to compare the performance of DIEG and PEG. The results show that DIEG has a much larger convergence range than PEG and has fuel optimality similar to that of the optimization method in most fault scenarios. Finally, the robustness of DIEG under various deviations is verified through Monte Carlo simulation.

Two-layer fault diagnosis model of aircraft based on LSTM

Different component malfunctions of the aircraft may cause fault characteristics crossing easily and similar trajectory deviations, making traditional single-type diagnostic approaches to risk misidentification. To accurately ascertain the fault status based on sequential data such as flight trajectories and attitudes, this paper introduces a two-tiered fault diagnosis model leveraging long short-term memory (LSTM) neural networks. The first layer of the model discerns the fault mechanism, while the second layer pinpoints the specific fault type. Datasets encompassing four distinct actuator faults and four varying degrees of thrust drop faults (20%, 40%, 60%, and 80%,) are constructed for model training and validation. Simulation results show that employing a two-layer LSTM network enhances both training convergence and testing precision, achieving a diagnostic accuracy of 99.38%, surpassing the 96.74% accuracy of a single-layer network. Furthermore, the LSTM network's diagnostic accuracy, utilizing sequence information, significantly outperforms that of the Least Squares Classifier and the BP neural network, which rely on individual state information. Thus, the efficacy of proposed two-tiered LSTM-based fault diagnosis model in improving diagnostic accuracy and reliability in aircraft systems has been validated.

Angle of Attack Characteristics of Full-Active and Semi-Active Flapping Foil Propulsors

As a propulsor with a good application prospect, the flapping foil has been a hot research topic in the past decade. Although the research results of flapping foils have been very abundant, the performance-influencing mechanism of flapping foils is still not perfect, and the research considering three-dimensional (3D) effects for engineering applications is still very limited. Based on the above considerations, a systematic and parametric analysis of a small aspect ratio flapping foil is conducted to correlate the influencing factors including angle of attack (AoA) characteristics and wake vortex on the propulsive efficiency. Three-dimensional numerical analyses of full-active and semi-active flapping foils are carried out in this paper, in which the former focuses on different heave amplitudes and pitch amplitudes, and the latter concentrates on different spring stiffnesses. The analysis covers the full range of advance coefficient, which starts around 0 and ends at a thrust drop of 0. Firstly, the influence of the maximum AoA (αmax) on the efficiency and thrust coefficient of these two kinds of flapping foils is analyzed. The results show that for the small aspect ratio flapping foil in this paper, regardless of the full-active or semi-active form, the peak efficiency as high as 75% for both generally appears around αmax = 0.2 rad, while the peak thrust coefficient of 0.5 occurs near αmax = 0.3 rad. Then, by analyzing the wake flow field, it is found that the lower efficiency of larger αmax working points is mainly due to the larger vortex dissipation loss, while the lower efficiency of smaller αmax working points is mainly due to the larger friction loss of the foil surface. Furthermore, the plumpness of different AoA curves is compared and analyzed. It was found that, unlike the results of full-active flapping foils, the shape of the AoA curve of semi-active flapping foils with different spring stiffnesses is similar, and the relationship with efficiency is not strictly corresponding. This study is expected to provide guidance on both academics and industries in relevant fields.

Real-time fault-tolerant guidance for launch vehicle ascending flight under thrust drop failure

This paper presents a real-time fault-tolerant guidance method, which solves the autonomous rescue problem in the event of thrust drop failure during the final stage flight of launch vehicles. The core of the method is the online and optimal reconstruction of both the degraded target orbit and the subsequent flight trajectory. Aiming at achieving satisfied onboard computing performance and making the best use of the launcher’s residual energy, a unique relaxation-penalization model transformation method for the nonlinear orbit injection conditions is proposed, and a successive convexification algorithm is developed to solve the problem. Furthermore, both the two classic thrust drop failure situations, that is the mass flow rate drop and the specific impulse drop problems, can be resolved by the proposed algorithm. The average run-time of this orbit-trajectory joint optimization algorithm is below 150ms in an embedded ARM processor @1.5 GHz, which is adequate for onboard use. Using the above algorithm as a central infrastructure, a receding-horizon-strategy-based fault-tolerant guidance method is proposed. By performing the optimization repeatedly at a high frequency, the effects of parameter deviations and external disturbances can be absorbed for the most part, for example, the guidance algorithm is verified to be capable of tolerating at least ±5% thrust deviations. Comprehensive numerical and hardware-in-the-loop simulations are performed to demonstrate the computational performance, accuracy, and reliability of the proposed algorithms.

A novel in-flight thrust modulation technique for SPRM: proof of concept

Thrust modulation (termination or reduction) is one of the most sophisticated topics in rocket motors. Thrust termination is employed when combustion must be terminated to ensure proper stage separation, to avoid motor explosion, to attain certain range, or for research purposes. While liquid and hybrid propellant rocket motors have the ability to extinguish thrust, thrust termination in solid propellant rocket motors (SPRMs) need the use of specialized technology. In-flight thrust termination of (SPRMs) can be achieved by sudden and sufficient increase in throat area. Increasing throat area causes high depressurization rate which consequently leads to extinguishing the rocket motor. In contrast, in-flight thrust reduction can be theoretically attained either by reducing chamber pressure to a new equilibrium level or by reducing the divergent section length. The objective of the present paper is to prove the concept of using Linear Shaped Charge (LSC) for thrust modulation; a technique novel to SPRM applications. Thrust termination concept is proved via a set of static firing tests on a standard test motor. In addition, a mathematical model is developed to predict the drop in thrust due to cutting the nozzle at any arbitrary location. The model is validated by comparing with published experiments for thrust termination and reduction. A parametric study is conducted based on model results. It is confirmed that by cutting the nozzle along its divergent section, thrust reduction less than 20% can be attained. Further reduction can be achieved by cutting the nozzle along its convergent section.

Propulsion characteristics of self-pitching flapping foil

As a promising form of oscillating hydrofoil propulsion, self-pitching flapping foil (SPFF) has been widely used in ocean engineering applications such as ships, underwater vehicles and ocean energy devices. Due to the complexity of its mass spring system, it is still a lack of comprehensive parametric analysis to conclude the theoretical maximum efficiency of SPFF up to now. Particular interest of this work is centered on the influence of frequency ratio (r), spring stiffness ratio (K′), advance coefficient (J) and pitch axis (c0/c) on the propulsive performance of self-pitching flapping foil (SPFF), including induced propulsion force, propulsive efficiency and related wake structure. The analysis covers the range of full advance coefficient, which starts around 0 and ends at a thrust drop of 0. The influence of stiffness ratio (K’) on the SPFF performance was discussed in the large range of 0.1–100. Special attention is also paid to the impact of system resonance on performance. By employing an appropriate combination of various parameters, the highest efficiency of SPFF could reach 87.6%, which shows it could perform a satisfactory propulsion performance as a marine propulsion. This study is expected to provide guidance on both academics and industries in relevant fields.

Numerical study on the dynamic process of ramjet/scramjet mode transition in the integrated RBCC full flow path through multistage fuel injection strategies

The rocket-based combined-cycle (RBCC) engine usually completes the transition process from ramjet to scramjet mode under the Ma∞ = 5.5–6.0. The dual-mode (ramjet/scramjet) transient process is an important technology for engine operation. CFD modeling based on Reynolds average Navier-Stokes was for studying the influence of fuel injection position, high-temperature rocket jet and fuel throttling time on the combustion flow structure and transient characteristics of RBCC engine during the mode transiton . The results showed that whether RBCC engine could successfully achieve stable mode transition was mainly influenced by flame stability in the combustor. When the rocket shuts down, it was more suitable to choose the fuel injection arrangement which distributed downstream of the combustor to realize relatively smooth mode transition. The analysis shows that in the mode transition process, using the synergistic combustion-supporting effect of the cavity and the strut is beneficial to improve the fuel combustion efficiency and reduce the thrust fluctuation. Although the high temperature rocket jet with low mass flow rate cannot maintain the fuel combustion in the isolator at high Mach number, it can expand the high temperature combustion area width in the central flow of RBCC engine and increase the secondary fuel combustion efficiency. Therefore, keeping the rocket running at low mass flow rate can reduce the thrust drop caused by insufficient fuel heat release. When adopting a mode transition strategy that throttles the fuel linearly over time, the engine can go through the thrust dip more smoothly, but this correspondingly increases the time required for the internal flow field to stabilize.

Design of thrust augmentation control schedule during mode transition for turbo-ramjet engine

A thrust augmentation control schedule was designed for turbo-ramjet engine during mode transition process to achieve high thrust performance. First, a mathematical model of turbo-ramjet engine was established, analyzing the reason for the existence of an insufficient thrust working range under conventional mode transition control schedule. On this basis, a thrust augmentation control schedule was devised for the thrust insufficient range considering the matching between the air inlet available airflow and the engine demand airflow. The control schedule could increase the flow capacity of the engine and reduce the overflow drag of air inlet by opening the ramjet bypass while keeping the turbofan components working at the maximum state, which could improve the maximum installation thrust performance of the engine and achieve a better air inlet/engine matching. By further designing the optimal regulation rule for the rear variable area bypass injector (RVABI), the installed thrust of the engine could be increased by more than 37.93% at most, while simultaneously reducing the fuel consumption by 0.16%. Finally, the turbo-ramjet engine was simulated along the isobaric flight path. The simulation results revealed that under conventional mode transition control schedule, the thrust insufficient occurred when Mach number is above 2.7. The designed thrust augmentation control schedule could effectively improve the maximum thrust of the thrust drop zone and provide sufficient thrust margin for engine during the whole working process. At the same time, considering the smooth thrust and thrust augmentation requirements, a compromise mode transition control schedule between the thrust augmentation control schedule and the conventional mode transition schedule is proposed. This control schedule could address the problem of insufficient thrust at high Mach numbers while the smoothness transition of engine total thrust could be guaranteed.

Fault Detection and Diagnosis for Thrust Drop of Launch Vehicles Against Disturbances

Rapid detection and accurate diagnosis for thrust drop fault are still open problems for launch vehicles. Existing studies have achieved fruitful results in fault detection and diagnosis (FDD) methods, leaving the complex disturbances not fully considered. The disturbances degrade FDD performance and are very hard to identify. In this paper, an FDD method based on multiple deep neural networks is proposed, which can deal with the complex disturbances. First, we split the FDD into two parts to reduce complexity. In detail, a fully connected neural network is selected to detect fault and diagnose fault degree, and a long short-term memory neural network is carefully tuned to diagnose fault mode. An analysis of the relationship between flight states and fault conditions is employed to design input of the fully connected neural network and the long short-term memory neural network. Then, a novel prediction and compensation method is proposed to deal with the unobservable and time-varying disturbances. Without extra information, data from the inertial measurement unit and navigation system are just enough to eliminate the disturbances, whereas existing methods may need more sensors. At last, illustrative simulations demonstrate the effectiveness and correctness of the proposed method.

Equilibrium flight theory-based orbit entry capability evaluation and guidance method of launch vehicles with thrust faults

<p indent="0mm">Considering the nonfatal engine faults (e.g., abnormal thrust) of a launch vehicle in the outer atmosphere flight phase, a method based on equilibrium flight theory for the evaluation of the vehicle’s orbit entry capability and continuous thrust guidance is proposed. First, the mechanism of mission failure caused by abnormal thrust fault is analyzed from the perspective of flight dynamics; on this basis, the simplified model of launch vehicle flight dynamics is established in the LVLH coordinate system, and the approximate analytical theory of equilibrium flight is provided. The conditions of an “equilibrium flight” and a “quasi-equilibrium flight” are further analyzed, and the differences between their guidance laws are compared. Next, the method and process of the online evaluation and decision-making of launch vehicles with an abnormal thrust are listed. Finally, numerical simulation is performed to reveal the relationship between thrust drop and orbit entry capability, as well as verify the correctness and practicability of the analytical theory of an equilibrium flight comprehensively. This simple and effective method integrates online capability evaluation and optimal orbit guidance, which can provide theoretical and technical support for rapid autonomous fault handling of launch vehicles with thrust faults.

Autonomous Attitude Reconstruction Analysis for Propulsion System with Typical Thrust Drop Fault

The propulsion system is one of the important and vulnerable sub-systems in a strap-on launch vehicle. Among different failure modes, the thrust drop fault is the most common and remediable one. It degrades vehicle attitude tracking ability directly. To this end, this paper focuses on the design and application of attitude reconstruction problems with a thrust loss fault during the ascending flight phase. We firstly analyze the special failure modes and impacts on the propulsion system through a Failure Modes and Effects Analysis (FMEA). Then, six degrees of freedom dynamic and kinematic models are formulated, which are integrated into the Matlab/Simulink environment afterward. The above models’ validation is realized through numerical simulations with different fault severity. Simulation results show that the max attitude deviation is only 0.67° approximately in the pitch angle channel under normal conditions, and the flight attitude angle deviation is directly proportional to the thrust loss percentage when the thrust drop fault occurs. Based on the validated models, a practical reconfigurable ideal through adjusting the control allocation matrix is analyzed. Then, an automation redistribution mechanism based on the moment equivalent principle before and after the thrust drop is proposed to realize proportional allocation of virtual control command among the actuators. The effectiveness of the designed attitude reconstruction method is demonstrated through numerical simulations and comparison analysis under various fault scenarios. The results show that the rocket attitude can be quickly adjusted to the predetermined program angle within about 2.5 s after the shutdown failure of a single engine, and the flight speed and altitude can also reach the required value with another 17 s engine operation. Therefore, the designed control reconfiguration strategy can deal with the thrust loss fault with high practicability and can be applied to real-time FTC systems. Last but not least, conclusions and prospects are presented to inspire researchers with further exploration in this field.

Powered-Coast-Powered Guidance Reconfiguration Method of Launch Vehicle with Thrust Drop Fault

This paper describes the study of a powered-coast-powered guidance reconfiguration (PCPGR) method used to solve the autonomous rescue problem for the mission profile of a launch vehicle with a coasting phase when the thrust drop fault occurs in the first powered phase (FPP) of the final stage. We first described the constraints of the final stage and the construction of the PCP guidance problem. Then we evaluated the adaptability of the guidance reconfiguration (GR) offline with numerical optimization by adjusting the constraints of the FPP, the coast phase, and the second powered phase. To determine the fault state set where the rocket can enter the prescribed target orbit through the GR initiate, we proposed a Newton method-based rapid replanning method of the transfer orbit that transforms the complex multi-flight phase trajectory planning problem into a feasible transfer orbit search problem to produce a fast solution onboard. Combined with the adaptive adjustment of the coasting time and the iterative guidance mode, we realized the autonomous online rescue of the payload. The simulation results showed that the proposed method achieved a reliable and rapid solution and improved a launch vehicle’s adaptability to a thrust drop fault.

Mission reconstruction for launch vehicles under thrust drop faults based on deep neural networks with asymmetric loss functions

Launch failures of launch vehicles due to thrust drop faults have occurred throughout aerospace history, thus studying Mission Reconstruction (MRC) problems is crucial to alleviate the losses resulting from the thrust drop faults. Unlike the traditional trajectory optimization problem with a specified target orbit, the rescue orbit and flight trajectory in the MRC problems need to be optimized simultaneously. To solve the MRC problems rapidly and accurately, a deep neural network-based adaptive collocation method (DNN-based ACM) is proposed. In the offline part, using the datasets generated by convex optimization and adaptive collocation method, deep neural networks (DNN) are trained to map the relationship from the fault states to the optimal rescue orbital elements and terminal control variables. During the DNNs training, an asymmetric loss function is proposed to avoid the positive error of the semi-major axis which renders the trajectory optimization unsolvable. In the online application of the DNNs, the optimal rescue orbit is determined and proper initial guesses are set for the MRC problems. The flight trajectory from the fault position to the optimal rescue orbit can be effectively solved by the ACM with the proper initial guesses. Simulation results validate the effectiveness of the DNN-based ACM, and substantiate the high efficiency performance and optimal performance of the presented MRC method.

Improved Parallel-Structured Newton-Type Guidance for Launch Vehicles Under Thrust Drop Fault

When launch vehicles encounter thrust drop f4aults during endoatmospheric flight, conventional nominal trajectory-based guidance methods may fail to compensate adequately; and the launch mission may culminate in failure. In this paper, we present an improved parallel-structured Newton-type guidance (PSNG) algorithm to address the aforementioned challenge. Compared to the basic PSNG algorithm, the proposed algorithm entails two major improvements: determination of the free final time and handling of the path constraint using the projected Newton-type method. Except for inheriting the benefits of the original PSNG algorithm, the improved algorithm can solve the free-final-time and path-constrained ascent guidance problems directly. Numerical simulations with various ascent scenarios of thrust drop faults are conducted, and the results show that the improved algorithm has high computational efficiency, accuracy, and robustness against disturbances and uncertainties.

ACOUSTIC METHOD AND HARDWARE-SOFTWARE COMPLEX OF OPERATIONAL DIAGNOSTICS PRE-STAGE STATE OF GAS TURBINES ENGINES BASED ON INVARIANTS RANDOM PROCESS

One of the reasons for the early decommissioning of gas turbine engines (before the specified resources and service life are depleted) is the occurrence of unstable operating modes, namely surge. Surge is a violation of the stable operation of gas turbine engines, due to the occurrence of longitudinal self-oscillations of the flow in the entire airgas path of the gas turbine engine, resulting from the loss of dynamic stability of the flow, accompanied by a sharp drop in thrust and powerful vibration that can destroy the engine. An acoustic method for diagnosing the presurge state of gas turbine engines has been developed, based on an assessment of the change in the distribution of the amplitudes of a vibroacoustic signal, which is characterized by the use of an invariant characterizing the normal distribution of random variables, which makes it possible to quickly assess the presurge state regardless of the type, size and history of engine operation, which makes this method universal. The use of the developed acoustic method for determining the presurge state of gas turbine engines makes it possible to avoid the need for noise filtering, which at the present stage of development of presurge condition diagnostics systems based on vibration analysis plays a huge role. To implement the developed acoustic method, a hardware-software complex was created. The antisurge system can be used in automatic mode by installing hazardous and critically dangerous zones in the program for informative parameters and invariants characterizing the occurrence of surge.

Effect of fuel grain configuration on the thrust of a solid-fuel scramjet

The thrust of a solid-fuel scramjet rapidly decreases during the combustion process, limiting its application in hypersonic aircrafts. The purpose of this study was to alleviate this problem by optimizing the structure of the fuel grain. A 2D axisymmetric model was established to investigate the internal flow structure of the combustion chamber. Two types of fuel grain structures were proposed to reduce the thrust drop of the solid-fuel scramjet. A series of experiments were conducted to verify the correctness of the numerical model and the effectiveness of the fuel grain structures in suppressing the thrust drop. The results showed that the new fuel grain structures of the solid-fuel scramjet can slow down the total pressure loss in the combustion chamber and thereby suppress the thrust reduction.

High-Fidelity Modeling of Multirotor Aerodynamic Interactions for Aircraft Design

Electric aircraft technology has enabled the use of multiple rotors in novel concepts for urban air mobility. However, multirotor configurations introduce strong aerodynamic and aeroacoustic interactions that are not captured through conventional aircraft design tools. This paper explores the capability of the viscous vortex particle method (VPM) to model multirotor aerodynamic interactions at a computational cost suitable for conceptual design. A VPM-based rotor model is introduced along with recommendations for numerical stability and computational efficiency. Validation of the individual rotor is presented in both hovering and forward-flight configurations at low, moderate, and high Reynolds numbers. Hovering multirotor predictions are compared with experimental measurements, evidencing the suitability of the proposed model to capture the thrust drop and unsteady loading produced by rotor-on-rotor interactions.

Joint dynamic optimization of the target orbit and flight trajectory of a launch vehicle based on state-triggered indices

This paper solves the problem of the on-line joint optimization of the target orbit and flight trajectory of a launch vehicle experiencing a thrust drop failure, and proposes state-triggered indices to improve the solving efficiency. The ‘highest-height index’ is triggered first to find the maximum-height circular orbit (MCO) in the orbital plane formed at the time of failure. If the orbital altitude is less than the perigee height of the prescribed target orbit, the MCO will be regarded as the optimal circular orbit to ensure safety. If not, the ‘minimal orbital plane deviations index’ is triggered to subsequently find the optimal elliptical orbit (OEO), where the orbital inclination and longitude of the ascending node are preferentially adjusted on the basis of the MCO. Furthermore, if the OEO is coplanar with the target orbit, a ‘minimal orbital shape deviations index’ is triggered to find the optimal rescue orbit, which eliminates the deviations of the perigee altitude, semi-major axis and perigee argument on the basis of the OEO. Estimation of the geocentric angle of the injection point, construction of the approximated sub-problem, and convexification processing are adopted to provide initial guesses to boost the computing. Monte Carlo simulations show that the method not only converges to the optimal or sub-optimal solution for different failure states, which avoids suspended calculations, non-convergence, and constraint violations that are common when directly solving this type of rescue problem, but also meets the basic requirements of on-line applications.

Mathematical model and vector control of a six-phase linear induction motor with the dynamic end effect

This study investigates a short primary double-side six-phase linear induction motor (LIM) operating in non-periodic transient conditions. The main purpose is to solve the problems arising from the dynamic end effect. The equivalent circuit method is used for intensively studying the transient time-varying law of eddy current and excitation inductance, and the finite element method is adopted for verification. The calculation of eddy current loss leads to an expression of secondary equivalent resistance. Then, a mathematical model is established for the six-phase LIM involving the dynamic end effect. Furthermore, the matrix reconstruction method is adopted for the coupling of windings. Based on the time-varying characteristics of the model, an improved vector control method that can update the velocity-related correction coefficients through on-line calculation is proposed. During the transient process, the slip frequency and the torque current are modified in real-time to achieve the maximum output thrust of the motor. This approach can compensate for the thrust drop caused by the dynamic end effect, thereby improving the control precision. Simulations and experiments show that the related theory and the proposed control method are feasible and effective.

Autonomous trajectory planning for launch vehicle under thrust drop failure

An online autonomous rescue strategy and the algorithm for launch vehicle are studied in the case of a thrust drop during ascending flight. First, an iterative guidance method and numerical integration are used to evaluate the remaining carrying capacity. When the target orbit is unreachable, but the lowest safe orbit could be met, an optimal rescue orbit (ORO) needs to be solved. Using the geocentric angle estimation, orbit coordinate system transformation, and convexification, the maximum-height circular orbit in the orbit plane determined by states at failure time, is found under the constraint of geocentric angle. Then, an optimal circular orbit (OCO) relaxing the above constraint is solved by the adaptive collocation method (ACM) using the former solution as initial values. Finally, a decision whether to adjust other orbital elements by comparing the height between OCO and target orbit is made. The algorithm gives priority to the orbit height, optimizes the ORO based on the objective function weights, and makes full use of the advantages of geocentric angle estimation: simplifying the terminal constraints, making the convex optimization achieve good convergence, and improving the effectiveness of the ACM under reasonable initial guess. The simulation results show that the proposed strategy and algorithm are adaptable and, convergent for onboard application.