Journal of Aerospace Engineering

The effects of the shock wave and the expansion fan on the flat plate film cooling are numerically studied under the supersonic condition with inclined cylindrical film holes. A triangular-shaped shock generator is applied, with both the shock wave and the expansion fan induced. Numerical investigations are conducted under four shock generator wedge angles ranging from 0° to 10° and six Mach numbers of the coolant injection ranging from 0.6 to 1.6. The flow features induced by both the shock generator and the coolant injection as well as their effects are discussed. The film cooling effectiveness is compared between different coolant injection Mach numbers and shock generator wedge angles. The cooling jet can partially resist the interference of the shock waves. With the increase of the coolant Mach number, the negative effect of the shock impingement becomes smaller; however, the positive effect of the expansion fan is less affected. Consequently, with the increase of the shock generator wedge angle, the film cooling performance tends to decrease under a smaller jet Mach number, but it tends to increases under a larger jet Mach number. With the increase of the shock generator wedge angle, the optimal coolant injection Mach number for the best cooling performance gradually increases from 1.0 to 1.2.

This study conducted wind tunnel tests to verify a passive jet flow control method for suppressing the aerodynamic force on a chamfered square cylinder at a moderate Reynolds number, focusing on the influence of the opening parameters and the spanwise spacing ratio of hollow pipes on the jet flow control effect. The pressure coefficient distribution on the surface of the cylinder and the overall aerodynamic coefficient were considered in order to systematically analyze the influence of opening shape, opening number, opening height, pipe height, and pipe radial thickness, and to determine the optimal combination of opening parameters. Subsequently, we analyzed the influence of the spanwise spacing ratio on passive jet flow control. Furthermore, the optimal arrangement of the hollow pipe was derived by comparing the aerodynamic coefficients of uncontrolled and controlled cylinders. The results revealed that the passive jet hollow pipe significantly reduces the root mean-square value of the pressure coefficient and the fluctuation of the lift coefficient. Moreover, the platform area of the mean pressure coefficient at the leeward side of the cylinder increases and the mean drag coefficient decreases. The spectrum analysis results showed that the passive jet flow control pipe alters the frequency and intensity of vortex shedding in the wake region. All results indicated that the aerodynamic forces and wakes of a chamfered square cylinder are effectively controlled through jet flow control with perforated hollow pipes.

Abstract

Traditional aerodynamic modeling methods have some difficulties in multiparameter unsteady state-space modeling and control law design, which raise significant challenges in flight control. The rapid development of machine learning provides a new idea for unsteady aerodynamic modeling and control law design, which has significant value for theoretical research and engineering application. In this paper, an unsteady aerodynamic model environment was established based on deep neural network (DNN), and a dynamic computational fluid dynamics (CFD) virtual environment was taken as an approximation of the real environment. A deep reinforcement learning (DRL) longitudinal-axis flight control method was studied on the basis of both environments. The deep deterministic policy gradient (DDPG) algorithm was used to implement flight control, and the effects of different constraints in the reward function on the results of DRL were studied at the same time. The results show that the DDPG agent effectively achieves longitudinal-axis flight control in the model environment. The DDPG agent trained in the model environment was also used for longitudinal-axis flight control in the dynamic CFD virtual environment, and this method provides a reference for the coupling of dynamic CFD and DRL. The control results from the model environment and the dynamic CFD virtual environment are compared, and results show that the DDPG agent has good robustness in both environments.

Practical Applications

The results of this study suggest that machine learning and deep reinforcement learning can effectively solve complicated multiconstraint, uncertain, and sophisticated control tasks. In this paper, a DNN-based unsteady aerodynamic modeling method was established based on dynamic computational fluid dynamics calculation and machine learning, and the results show that the DNN-based unsteady aerodynamic model has advantages in dealing with highly nonlinear data under unsteady conditions. In addition, the deep deterministic policy gradient algorithm can be well applied in longitudinal-axis flight control, and the agent can be trained in DNN-based unsteady aerodynamic model environment. The agent was also used for longitudinal-axis flight control in the dynamic computational fluid dynamics virtual environment, and the results show that the agent has good robustness in both environments. This provides the possibility for the agent to achieve longitudinal-axis flight control in a real environment.

A propeller blade has a significant effect on the performance of a propulsion system. It is of great significance to develop an efficient and accurate optimization strategy for propeller design in the preliminary stage. A key concept of the design is to establish a fast hybrid aerodynamic optimization framework employing the reformulated Blade Element Momentum (rBEM) model to optimize the propeller blade shape under cruising conditions. First, the rBEM numerical model is established and validated with the experimental data in both hover and forward-flight cases. Second, a fast hybrid design optimization strategy based on the rBEM model is developed, which divides the propeller design process into gradient-free planform optimization and gradient-based airfoil optimization. The planform and sectional designs are coupled with each other in an iterative manner until convergence criteria is reached. Based on the proposed optimization strategy, the hybrid design process shows gradual convergence. The efficiency of the optimized propeller has been increased by 29.84% when compared with the baseline propeller. The solidity ratio of the optimal propeller is smaller than that of the baseline propeller, which has a smaller chord length distribution, a larger twist angle distribution, and lesser camber and thickness of the cross-sectional airfoil.

Abstract

Accurate methods for geometrically exact aeroelastic stability analysis of composite helicopter rotor blades in hover have been proposed in this paper. The aeroelastic model is established by using the geometrically exact beam theory and the updated variational asymptotic beam sectional (VABS) analysis which can cover the effect of initial twist and curvatures and calculate accurately the structural properties of the blade cross section with arbitrary shape and material distribution. The Peters finite state airloads theory and Peters-He finite state dynamic wake model are adopted to calculate the three-dimensional unsteady airloads. To ensure the calculation accuracy of the geometrically exact aeroelastic stability of the blades, the time domain method based on the finite element spatial discretization, Newmark numerical integration and Newton-Raphson methods to calculate the aeroelastic responses of blades and the moving-block analysis to extract the regressive lag mode damping from the blade transient responses are established. The accuracy of the proposed methods is verified by experimental results. The investigation results indicate that the transverse shear deformation and initial curvatures of the blades have significant effects on the aeroelastic stability of the hingeless composite rotors in hover.

Practical Applications

The rotor blades of helicopter are the lift components of helicopter. Under the flight conditions of high speed, heavy load, and large maneuvering, the composite blades will produce large aeroelastic deformation and aeroelastic instability. Therefore, it is necessary to establish an accurate aeroelastic analysis method for composite blades of helicopters. In this paper, an accurate method for geometrically exact aeroelastic stability analysis of composite rotor blades of a helicopter in hover has been proposed with applications to sophisticatedly treat the anisotropy material properties, arbitrary cross-sectional shape and material distributions, and initial twist, and material distributions and initial twist and curvatures of composite rotor blades of a helicopter. The cross-sectional structural properties and large deflections of the blades under applied forces and moments can be calculated accurately and efficiently. The presented method may also be applied to the analysis of turbine blades, aircraft wings, and many other engineering structures with beam-like geometries and has potential applications in the structural and aeroelastic modeling of these beam-like structures.

This paper considers the problem of unmanned aerial vehicle (UAV) aerial collision net recovery in a three-dimensional non-decoupling environment and approach angle constraint. A robust control-observer framework-based guidance law was designed via the nonsingular fast terminal sliding mode control (NFTSMC) technique and adaptive sliding mode disturbance observer (ASMDO). To estimate the disturbance of the guidance system in finite time, an ASMDO is presented in which the parameters are autonomously adjustable according to the estimation error. The proposed control implementation uses the nonsingular fast terminal sliding mode (NFTSM) technique to drive the line-of-sight (LOS) angle error and LOS angular rate fast convergence under model coupling and external disturbance. Furthermore, regarding the actuator of UAV with second-order dynamic, a backstepping guidance law to compensate for actuator dynamics is proposed with the aid of a finite-time converged differentiator, which can estimate directly the derivative of the virtual control law and guarantee the finite-time convergent characteristic of the partially integrated guidance and control system. Simulation studies and comparisons verified the efficiency of the proposed guidance law in the presence of a complex disturbance lump.

In this study, the whisker structure of a harbor seal is utilized on the trailing edge of a high-pressure turbine stage to control the interaction between the wake and the blade. The paper investigates two cases: a prototype high-pressure turbine stage, and a biomimetic turbine stage where the vane and blade trailing edges are modified to mimic the whisker structure. The flow in the prototype and the biomimetic turbine stage is simulated using the delayed detached-eddy simulation (DDES) method. Results show that fluid behind the whisker trailing edge converges to the valley region. The whisker trailing edge transforms the Karman vortices into hairpin vortices, narrowing the wake area of the vane. A narrower wake area delays the interaction between the wake core area and the blade. The prototype vane’s pressure-side and suction-side shedding vortices interact with the leading edge of the blade, resulting in large pressure fluctuations. The whisker trailing edge weakens the impact of the wake on the blade’s leading edge, significantly reducing the amplitude of pressure fluctuation at the blade’s leading edge. Meanwhile, the whisker trailing edge minimizes the adhesion of the small vortices on the blade, suppressing the high-frequency pressure pulsation at the blade’s leading edge. Moreover, the breaking of the Karman vortex suppresses the rotational velocity field in the blade passage. The suppression of the rotational velocity field of the shedding vortex reduces the velocity fluctuations in the blade’s near-wall region and boundary layer.

Wind tunnel experiments were conducted to explore the control effect of a synthetic jet actuator array placed at the leading edge of a National Advisory Committee for Aeronautics (NACA) 0015 airfoil at two Reynolds numbers of $1.1\times {10}^5$ and $1.6\times {10}^5$. Synthetic jets were generated perpendicular to the leading edge with high-frequency excitation to imitate the biomimetic effect of leading-edge tubercles. Force measurements over the angle of attack $\alpha$ from 0° to 40° display similar aerodynamic characteristics between the present synthetic-jet–controlled airfoils and previous leading-edge tubercled ones in both the prestall and poststall regimes. Thus, virtual tubercles constructed by synthetic jets can be used to effectively improve poststall aerodynamic performance of airfoils. Flow field characteristics at $\alpha =8°$, 16°, and 32° were analyzed to reveal the control mechanism at different angles of attack. At $\alpha =8°$, synthetic jets ruin the high-velocity flow region on the suction surface, resulting in attenuated aerodynamic performance in the prestall regime. In the poststall regimes, however, synthetic jets can directly inject high momentum into the separated shear layer at $\alpha =16°$, which energizes the boundary layer and thus delays flow separation. As $\alpha$ increases to 32°, synthetic jets can still improve the lift performance to some extent by enhancing momentum exchange between the outer high-velocity flow and inner low-velocity separation region. In particular, synthetic-jet tubercles can provide the potential for better practicality and higher efficiency than traditional leading-edge tubercles in improving unsteady aerodynamic characteristics.

A manipulator mounted on a free-floating satellite will be a key technology in active debris removal and on-orbit servicing missions. Such a system is characterized by a dynamic coupling between the manipulator and the satellite. Thus, there is a need for model-based control systems that take the satellite’s motion into account. As it is assumed that the satellite is uncontrolled during the manipulator’s operations, the satellite-manipulator system is nonholonomic, which makes a stability analysis complex. In addition, it is difficult to precisely identify mass and inertia of the system. To overcome these problems, a backstepping adaptive control system is proposed. Stability is proved using Lyapunov theory. There are three main novelties of the approach: (1) the backstepping algorithm is used for operational space control of a free-floating space manipulator; (2) adaptation laws are derived directly, using the full linearly parametrized dynamic model, with conserved, but nonzero initial momentum and angular momentum; and (3) satellite velocity measurements are used as the desired trajectory in order to allow application of the backstepping algorithm for the full dynamic model with nonzero initial momentum and angular momentum. Control quality is validated using Monte Carlo simulations. Our results are compared with a control scheme based on the kinematic Jacobian with satellite velocity feedback. Our adaptive control system not only performs better in terms of control quality, but also ensures asymptotic stability.

This paper proposes an analytic approach for impact time control guidance laws against stationary targets using biased proportional navigation. The proposed guidance scheme realizes the impact time control in two different ways: the first approach directly uses the exact time-to-go error to satisfy both the impact time control and the field-of-view constraint, while the second approach adopts a look angle tracking law to indirectly control the impact time, with the reference profile of the look angle generated using the exact time-to-go solution. The stability properties of the proposed guidance laws are discussed, and numerical simulations are carried out to evaluate their performance in terms of accuracy and efficiency.

A new bio-inspired flapping wing rotor (FWR) configuration combining the high lift of a flapping wing (FW) with the vertical take-off and landing characteristics of a rotor has proved its concept feasibility in the last decade and can be used for reconnaissance and aerial photography missions in complex and confined environments. During high-frequency flapping motion, the leading and trailing edge vortices of the flapping wing rotor alternately shed and form vortex sheets in the wake, resulting in periodic lift and a highly unsteady flow field. In the past, unsteady potential flow aerodynamic methods were derived based on a two-dimensional (2D) airfoil and extended to three dimensions by the blade element theory, neglecting the spanwise evolution effect of the vortices. In this paper, an unsteady vortex ring method (UVRM) model was developed based on two-dimensional (3D) potential flow theory for the aerodynamic calculation of a flapping wing rotor, which takes into account the spanwise evolution of shedding vortices at the leading and trailing edges of the flapping wing. The calculation results of the UVRM for several typical cases were compared with those of the computational fluid dynamics (CFD) method to verify the reliability of the UVRM. Then, fluid–structure interaction analysis for the flapping wing rotor was conducted based on the UVRM model, and the phase difference between twist deformation and flapping angle was found to be approximately 90°. In addition, the effects of wing chordwise and spanwise deformation on the aerodynamic performance of the flapping rotor were investigated based on the UVRM. The results show that the wing chordwise deformation has a positive effect on the mean lift coefficient of the flapping rotor. The wing spanwise deformation has a negative effect on the mean lift coefficient of the flapping rotor. The unsteady vortex ring method model proposed in this paper provides a solution to obtain the aerodynamic force of the flapping wing rotor in an efficient and accurate way.

The high-quality flow field of an aeroengine ground test cell is the basis for accurate and reliable engine testing and obtaining relevant performance data. In this study for a U-shaped ground test cell, first, the similarity dimension analysis between the original model and the $1:10$ scaled model is conducted to verify the accuracy of the numerical simulation of the scaled model. Then, the parametric numerical simulation method is used to study the effects of the augmentor’s diameter and the distance between the engine jail nozzle and the augmentor inlet on the aerodynamic flow field of the ground test cell for the $1:10$ scaled model. The results show that if the diameter of the augmentor tube is too large or too small, the flow rate of the bypass airflow in the augmenter will decrease; thus, the average temperature of the exit section will increase. At the same time, the change of airflow exhaust speed in the blast basket can effectively affect the uniformity of airflow in the silencing tower. With the increase of distances between the engine jail nozzle and the augmentor inlet, the ejection effect of the engine tail jet enhances, but the exhaust velocity in the blast basket decreases; accordingly, the quality of the flow field in the silencer tower significantly deteriorates with the tendency of the vortex centers on both sides of the pipe moving upward. In $\mathit{Case}4$, the airflow velocity at the front chamber inlet of the test workshop is significantly higher than that of the baseline model, $\mathit{Case}0$, and the ejection ratio increases by 19.4%.

Sandwich panels have attracted widespread attention owing to their lightweight and ability to meet strength, energy absorption, and thermal resistance requirements. Therefore, they are useful in the automotive, aerospace, marine, and infrastructure industries. This study investigated the aeroelastic behavior and vibration characteristics of foam-filled composite corrugated core sandwich panels at supersonic speed under different equal-mass conditions. The aerodynamic pressure was calculated using the quasi-steady first-order piston theory, and the aeroelastic motion equations of sandwich panels were established using Hamilton’s principle. The effects of the foam-filled composite corrugated core thickness on flutter dynamic pressures and the natural frequencies were analyzed when the mass of the foam-filled composite corrugated sandwich panel was $n(0<n\le 1)$ times that of the basic composite laminate panel. The results obtained indicate that the flutter dynamic pressures and the natural frequencies can be improved by changing the foam-filled composite corrugated core thickness while keeping the mass the same. Compared with the basic model, the mass can be reduced by more than 70% using a foam-filled composite corrugated panel at the same natural frequency, and the mass is reduced by more than 50% at the same flutter dynamic pressure. This is highly effective at reducing the aircraft panel weight.

Abstract

The current study experimentally investigates the influence of nonflat plate trailing edge serrations on airfoil noise characteristics. Experiments are carried out at different flow velocities in the range of $25–40\mathrm{m}/\mathrm{s}$, corresponding to the Reynolds number range of ${2.4\times 10}^5$ to ${3.9\times 10}^5$. The wavelength ($\lambda$) of the serrations considered is in the range of 5–30 mm, and the serration height ($2h$) is in the range of 10–30 mm. The acoustic spectra show that the conventional nonflat plate trailing edge serrations generate a narrowband vortex shedding noise. The narrowband noise is observed to decrease with an increase in the wavelength; however, it is observed to increase with an increase in the serration height or root thickness by around 16 dB. Inclinations ($\theta$) provided at the root of the serrations reduce the narrowband noise up to 20 dB and the broadband noise up to 5 dB. When a perforated plate inserts the sawtooth gap of the serration, the hybrid configuration effectively eliminates the vortex shedding phenomenon at the roots with a significant reduction in the low frequency range as well as the high frequency broadband noise.

Practical Applications

Airfoil noise poses a significant challenge in aviation, wind energy, and other sectors where aerodynamic noise reduction are crucial. When an airfoil is subjected to turbulent flow, the primary source of noise arises from the interaction of the flow with its trailing edge. The geometrical modifications made at the trailing edge helps to minimize the intensity of the noise generation. This paper presents the results of an experimental investigation on the impact of nonflat plate trailing edge serrations. The major findings of the study reveal that employing a nonflat plate trailing edge serrations can substantially mitigate airfoil trailing edge noise levels.

Low-thrust propulsion systems are widely used in space missions due to their high specific impulse and reduced propellant consumption. However, current research focuses on eliminating drifts in low-thrust maneuvers, resulting in high operational costs for daily station-keeping. This study aims to achieve autonomous station-keeping by exploiting long-term periodic behaviors caused by geostationary resonances. First, resonance maps caused by the Earth’s gravity and lunisolar perturbation are revisited. Then, station-keeping slots are selected based on a phase space study of the resonances. A closed station-keeping track is formed by designing a low-thrust arc to work with the drifting arc. This track moves along the edge of the station-keeping slot, providing a more extended control period and reducing ground-station operational costs. Finally, an indirect optimizing method and efficient initial costate guess technique are proposed for calculating low-thrust maneuvers.

The abort mission refers to the mission where the landing vehicle needs to terminate the landing mission when an anomaly happens and be safely guided to the desired orbit. This paper focuses on solving the time-optimal abort guidance (TOAG) problem in real-time via the feature-based learning method. First, according to the optimal control theory, the features are identified to represent the optimal solutions of TOAG using a few parameters. After that, a sufficiently large data set of time-optimal abort trajectories is generated offline by solving the TOAG problems with different initial conditions. Then, the features are extracted for all generated cases. To find the implicit relationships between the initial conditions and identified features, neural networks are constructed to map the relationships based on the generated data set. Finally, experimental flight tests are conducted to demonstrate the onboard computation capability and effectiveness of the proposed method.

The sequential approximate optimization (SAO) method is widely adopted in aircraft optimization design, whose further improvement in efficacy and efficiency has encountered bottlenecks. In this paper, a collaboratively iterative sequential approximate optimization (CISAO) method with field metamodel (FM) accelerating computational fluid dynamics (CFD) is proposed for aerodynamic optimization problems to enhance the coupling between metamodels and CFD high-fidelity models (HFMs). In this method, the field metamodel is used to predict the flow field distributions with respect to the design variables to accelerate the numerical solution of CFD HFMs. Two benchmark aerodynamic optimization cases are performed to validate the proposed method. The time costs for single CFD HFMs in the two test cases are shortened by 37.76% and 12.11%, respectively, and the total time costs of the two test cases are shrunk by 64.11% and 44.47% compared with the conventional SAO method. The optimization results indicate that the proposed method significantly reduces the time costs and improves the optimization efficiency. The proposed method shows a promising prospect in aircraft optimization design.

During the launch phase of a multinozzle rocket, the high-level noise induced by the supersonic jet ejected from the nozzles significantly affects the structural safety of the rocket and the reliability of the instruments. Therefore, it is important to predict and suppress the multinozzle rocket jet noise. This study proposes a numerical analysis method that uses the detached eddy simulation model combined with species transport equations, the discrete-phase model, and the acoustic analogy method (FW-H). It aims at a four-nozzle liquid rocket to predict jet and noise environments while considering the physical effect of evaporation and the influence of the launch pad structure. The proposed method was verified by comparing the single-nozzle jet simulation and experiment results. The prediction of the supersonic jet flow field and noise environment at different lift-off heights of the rocket was completed. An annular water jet noise-reduction system was designed to suppress the noise environment at different heights by single-deck and double-deck water injections. The noise attenuation effect of the water injection system was better when the lift-off altitude increased. And the noise suppression effect of the double-deck water injection system was more comprehensive than that of the single-deck water injection system. Furthermore, the reasons for the difference in noise reduction by different layers of water injection were analyzed in detail.

This paper proposes a reinforcement learning–based adaptive attitude control (RLAC) method for a class of hypersonic flight vehicles (HFVs) output constrained nonaffine attitude control problems subject to unmatched disturbances. First, by considering the strong coupling of HFVs attitude dynamics, the uncertainty of aerodynamic parameters and the complexity of the flight environment, a second-order multivariable nonaffine nonlinear control system is obtained. Then, by introducing specific nonlinear function and coordinate transformation techniques, the output constrained nonaffine control problem is transformed into a stabilization problem of several new variables. Moreover, dual actor-critic networks and their adaptive weight update laws are designed to cope with unknown unmatched and matched structural uncertainties. Meanwhile, two super-twisting disturbance observers integrated with dual actor-critic networks are designed to compensate unknown unmatched and matched external disturbances. With the help of the Lyapunov direct method, output constraint, convergence of the estimated weights, and stability of the system are proved. Finally, the validity as well as superiority of the proposed method are verified by numerical simulations.