Journal of Aerospace Engineering

For the canard rotor/wing (CRW) aircraft, the elliptical wing with blunt trailing edge poses challenges to control surfaces and flap design. Based on the characteristics of CRW aircraft, an active circulation control (CC) system was introduced for aircraft control and lift enhancement. The control system was designed and investigated based on numerical simulation. A benchmark problem was used to validate and calibrate the numerical simulation method of circulation control. Design parameters of two-dimensional (2D) circulation control system was determined by variant parameter analysis. A circulation control curved surface was optimized based on the above design parameters, and a well-performed 2D flow control system was achieved. Finally, the 2D design configuration was scaled and lofted to generate the three-dimensional (3D) circulation control wing (CCW). The 3D flow control wing was evaluated and analyzed. The results showed that the flow control concept was effective and can meet the lift enhancement and manipulation requirements of the CRW aircraft.

The completed study investigated the creation of Martian regolith-simulant composites. The goal was to see whether the type of Martian regolith simulant would influence the Martian regolith-simulant composite parts’ mechanical property values. Test samples of Martian regolith-simulant composites were created by combining four different types of Martian regolith simulant at different weight-loading percentages with polypropylene via twin screw extrusion and injection molding. Sample parts underwent tensile, flexural, Izod impact, dynamic mechanical analysis, and heat deflection temperature tests according to American Society for Testing and Materials standards. Two-factor analysis of variance tests were also completed on the averaged numerical data. The test results collected suggested that Martian regolith-simulant type (or source location) impacted a few of the mechanical properties of the Martian regolith-simulant composite sample parts created, including tensile modulus, flexural strength, and glass transition temperature. The degree of influence the source location type had on tensile modulus and flexural strength for the Martian regolith-simulant composites created was inferred to be low, as the $F$ values found were relatively close to the cutoff $F$ values. The analysis of variance test results also provided evidence that weight-loading percentage impacted all of the found mechanical property values of the Martian regolith-simulant composite sample parts created. The degree of influence the weight-loading percentage had on all of the mechanical property values was suggested to be high, as the respective $F$ values were much higher than the cutoff $F$ values. Regarding material processing, the Martian regolith simulant designated Sulfate ISRU was found to have unanticipated expansion characteristics, making it challenging to create usable parts with injection molding. These findings were aimed at providing further insight into the potential of Martian regolith-simulant composites.

High-lift airfoil slat and passive control noise were studied experimentally. A serrated slat with different parameters was installed on the trailing edge. Far-field noise was measured in an aeroacoustic wind tunnel using a microphone array. The results showed that the generation mechanism of discrete noise in slat noise was closely related to the feedback loop in the slat cove. As the Mach number increased, the pure tone frequency moved to the right, and the peak value gradually increased, whereas the pure tone frequency was not affected by the angle of attack. Serrated slats with different sawtooth angles had different effects on the aerodynamic noise of the slat, and the frequency of discrete noise also changed. The serrated slat needed to reach a certain value in the sawtooth angle to have the effect of noise reduction. When the sawtooth angle was large, the effect of noise increase also occurred, and there was a phenomenon of mode switching. The selection of the dominant mode was also determined by the time of modal occurrence. The smaller the sawtooth angle, the more obvious the effect of slat noise. When the sawtooth angle reached a certain angle, the discrete noise disappeared, and the noise reduction effect was basically the same. The serrated slat had little effect on the lift coefficient at a low angle of attack.

The closure parameters introduced in transition models can compromise the accuracy of prediction results. This paper investigated the uncertainty caused by the closure parameters in transition models ($\gamma$ and $k-\omega -\gamma$), and identified the key parameters that have the greatest impact on quantities of interest. The uncertainty was propagated by the point-collocation no-intrusive polynomial chaos method, and the relative contribution of each parameter to uncertainty was assessed by the Sobol index. The natural transition flow (Schubauer–Klebanoff flat plate) and low-velocity airfoil flow (NLF0416 airfoil) were chosen for computational cases. The results, which are highly dependent on the closure parameters, validate the necessity of uncertainty research. The parameters of the $k-\omega -\gamma$ model are more sensitive than those of the $\gamma$ model. In the $\gamma$ model, the most critical parameter is $C_{TU2}$, which is negatively related to the transition position, i.e., the transition delays when $C_{TU2}$ increases and advances when $C_{TU2}$ decreases. In the $k-\omega -\gamma$ model, the most critical parameter is $C_2$, which is positively related to the transition position.

Abstract

This paper focuses on the shock wave/boundary layer interaction (SWBLI) at the compression corner. The concept of dynamic microvortex generator (Dy-MVG) is proposed based on the flow control mechanism. A high-precision drive system is used to control the periodic oscillation of the microvortex generator, thereby injecting controllable dynamic disturbances into the boundary layer. Based on the Feng Dong (FD)-06 wind tunnel, the control effect and mechanism of Dy-MVG on SWBLI are investigated by schlieren and pulsating pressure test techniques. The operating conditions are as follows: the incoming Mach number is 2.5 and 3.5, the total temperature is 300 K, and the oscillation frequency of the Dy-MVG is $2–60\mathrm{Hz}$. The results show that classical MVG moves the shock foot position downstream and increases the shock angle. Dy-MVG enables dynamic changes in the shock foot position and shock angle over a greater range than classical MVG. Compared with classical MVG, Dy-MVG further weakens the shock intensity in the SWBLI region, and the shock intensity decreases with the increase of excitation frequency. Dy-MVG further increases the pulsating pressure and the energy of low-frequency characteristics in the SWBLI region, resulting in better control of the SWBLI.

Practical Applications

The Dy-MVG proposed in this paper has better performance than the classical MVG in controlling SWBLI. Due to the adjustable characteristics of Dy-MVG, it can be applied to a wide speed range of aircraft. For example, it can be used for airfoil flow control to improve airfoil characteristics, and it can also be applied to the inlet to improve the inlet performance. The frequency and amplitude can be controlled to meet the requirements of airfoil and inlet under different flight conditions.

Shock wave–turbulent boundary layer interaction (SWTBLI) exists widely in the internal and external flow fields of supersonic and hypersonic vehicles, and has a significant influence on the aerodynamic performance. Separation scaling analysis for SWTBLIs has been pursued by researchers for decades. This paper took into consideration the downstream expansion fan to present a generalized scaling analysis of the separation zones induced by three types of interactions—incident SWTBLIs, compression corner SWTBLIs, and normal SWTBLIs. To establish this scaling correlation, SWTBLI flow-field data under the condition of free-stream ${\mathrm{Ma}}_0$ of 3.5 and ${\mathrm{R}}_{\delta }$ of $2.34\times {10}^5$ were simulated numerically. The influence of the expansion fan on the wall pressure distributions and the scaling of the separation zone was analyzed in detail. The size of the separation zone was found to be related closely to the pressure increase at the reattachment point. The scaling of the separation zone and the reattachment pressure increase are rendered dimensionless by the displacement thickness of the incoming boundary layer and the dynamic pressure of the incoming free stream, respectively. The generalized scaling of the separation zone was found to follow a simple power function against the reattachment pressure coefficient. This generalized scaling correlation fits a wide range of published experimental data and numerical simulation results.

This paper focuses on the track-to-object association problem based on the two-line element (TLE) set. A five-step track-to-object association algorithm is designed for the radar observation data. First, for all objects in the TLE set, their orbits are propagated and stored with a step of 10 min. Then the orbital position at an arbitrary epoch can be obtained through the interpolation algorithms. Second, for tracklets longer than 60 s, the extended Laplace method is implemented to determine the orbital elements. Third, the semimajor axis, eccentricity, inclination, and the right ascension of the ascending node are used to filter the TLE sets. More than 90% of the objects can be filtered, improving the association efficiency significantly. Fourth, a first-order association process is implemented to get candidate objects, with the difference between the observed positions and the TLE orbit positions as a metric. Finally, considering the measurement error and the TLE orbit uncertainty, an uncertainty-based final association process checks the candidate objects and gives the final association result. The proposed approach is tested with simulated and observed data. The true-positive rate is higher than 99.5%, while the false-positive rate and the false-negative rate are both lower than 0.1%.

Abstract

Three-dimensional numerical simulations are conducted to investigate supersonic flow over capsule/canopy two-body configurations at a freestream Mach number of 2.0 and a freestream unit Reynolds number of $2.04\times {10}^7{\mathrm{m}}^{-1}$. Four cases representing different flow modes with varying distances from the forebody (capsule) to the afterbody (canopy) are carried out with detached eddy simulations. The flow evolutions and aerodynamic properties of the configurations are investigated. The pressure fluctuations on the inner surface of the afterbody show a strong correlation with the position of critical shear layers (containing high turbulence kinetic energy) in the wake of the forebody. A large pressure fluctuation is observed when the shear layer acts on the inner surface and vice versa. The drag of the afterbody mainly contributed by the pressure forces first decreases and then increases as the distance between the two bodies is increased. In addition, by analyzing the lateral force of the afterbody, it is found that the proximity of the bodies only has a weak effect on the lateral stability performance.

Practical Applications

The present study can provide suggestions for the design of supersonic parachutes. By investigating the flow evolution of four cases with varied proximity of the capsule/canopy two-body configurations ($X/d$), it is found that only a configuration with a large enough $X/d$ (Case D in the present study) can sustain a bow shock of the canopy, leading to a more stable flow dynamic in comparison with the other three cases. Besides, Case D also features good aerodynamic performance, as the case sustaining a bow shock can provide a large magnitude of inner-surface pressure of the canopy due to the subsisting upward step of pressure (The step occurs when fluids cross shocks). Thus, a well-design supersonic parachute should be characterized by a large enough $X/d$. Additionally, in Case D, there is a conical shear layer generated when the wakes of the capsule cross the bow shock of the canopy. The shear layer contains high turbulent kinetic energy, and a large pressure fluctuation is observed when the shear layer acts on the inner surface and vice versa. Therefore, “avoiding the shear layer acting on the inner surface of the canopy” may also be a useful suggestion for designs of supersonic parachutes.

Abstract

Space manipulators are usually used to assist or replace the astronaut to execute on-orbit tasks. However, due to the burdensome tasks and harsh space environment, free-swinging joint failure is prone to happen. Because the faulty joint cannot provide torque, the motion capability of the manipulator will degrade, easily causing the task to be unable to continue. To satisfy the task requirement, a motion capability optimization method is proposed. First, we establish the kinematics and dynamics model of the faulty manipulator. Second, the dexterity and load operation ability indexes of space manipulator with free-swinging joint failure (SMFSJF) are constructed. Third, by using a sixth-order polynomial function with superfluous parameter to describe the trajectory of the active joint and end-effector, a trajectory optimization model is established. Through solving the optimization model based on the particle swarm optimization (PSO) algorithm, the optimal trajectory of SMFSJF is obtained. The simulation experiments are carried out with a 7-degree of freedom (DOF) space manipulator, verifying the correctness and effectiveness of the optimization method. The method has two advantages: (1) the motion capability of space manipulators with free-swinging joint failure is analyzed for the first time; and it is pointed out that compared with the faultless manipulator, the motion capability of the faulty manipulator is greatly degraded; and (2) by constructing the trajectory optimization model with superfluous parameter, the trajectory optimization is realized based on the PSO algorithm. It ensures that the task can be executed successfully by the manipulator with the optimal motion capability.

Practical Applications

Space manipulators are usually used to assist or replace the astronaut to perform on-orbit tasks, such as spacecraft docking, satellite release, and space station maintenance and assembly. However, due to the long-term service and harsh space environment, free-swinging joint failure is prone to happen. This kind of space manipulator also belongs to an underactuated system, whose input dimensions are fewer than the output dimensions. Because the faulty joint cannot provide torque, the motion capability of the underactuated manipulator will degrade. So a motion capability optimization method is proposed in this work. We analyze the motion capability of this kind of space manipulator according to its kinematic and dynamic characteristic for the first time. We found that compared with the faultless manipulator, the motion capability of the underactuated manipulator is greatly degraded. The proposed optimization method ensures that the task can be executed successfully with motion capability optimizing. There may be underactuated joints existing in other kinds of space manipulator to lighten the weight of drive motor and reduce launch costs. So the trajectory optimization method can also be applied to such space manipulators.

Large drag force and severe heating are two major issues encountered in the development of hypersonic vehicles. As an active control technology, the application of a counterflowing jet to hypersonic vehicles to reduce the wave drag and heat flux has attracted extensive attention. However, the flow instability associated with the counterflowing jet, particularly the instability related to the large oscillation of bow shock in front of the vehicle at a long penetrating mode, remains to be further studied. In this study, a numerical investigation for a hemispherical nose with a counterflowing jet in a hypersonic flow is carried out by solving the Navier-Stokes equations, based on a solver developed on the OpenFOAM platform. The instability characteristics of bow shock, the drag reduction mechanism, and the flow structure around the nose specially for a long penetrating mode are studied in detail. The results show that with a counterflowing jet, the shock standoff distance can be increased largely and the drag force reduced significantly, and under a specific jet pressure ratio the averaged shock standoff distance and drag coefficient can reach to the values of 3.5 and 0.45 times those without a jet, respectively. The drag force varies periodically as the shock wave moves back and forth in front of the hemisphere and the oscillation frequencies of the force and the shock standoff distance are the same, but there is a phase difference between them. It is found that the bow shock oscillation at a long penetrating mode is usually asymmetrical in time, and the oscillation also occurs at a short penetrating mode when the jet pressure ratio is in a certain range, but it is symmetrical in time and the amplitude of the oscillation is usually smaller. The results of present study may provide useful information for the practical application of counterflowing jets.

As the goal of returning to the Moon and building a lunar colony becomes more prominent within the space industry, research and experiments in lunar exploration and establishment are steadily increasing. Among the many resources required for lunar-centric research, lunar simulant is one of the most vital, with multiple uses across various fields, such as engineering, chemistry, and biology. This study focuses on the development of the alternative Thailand lunar simulant 01 (TLS-01A), which utilizes raw material sourced from locally available mines in the Chanthaburi-Trat area in Thailand. This simulant was processed using simpler methods than the previously developed Thailand Lunar Simulant (TLS-01) and was geared towards small-scale production. First, production began with the selection of the basaltic rock raw materials sites. Then, we ground the raw material into finer sizes to replicate the lunar regolith. Next, a scanning electron microscope (SEM) was used to analyze grain morphology. Geochemical properties were determined using wavelength dispersive X-ray fluorescence (WDXRF) and inductively coupled plasma mass spectrometry (ICP-MS). Finally, direct shear tests were conducted to analyze the physical properties of the simulants. The processed simulant was found to exhibit rough and sharp edges, while its friction angle and cohesion value were 29° and 3.05 kPa, respectively, which are in line with well-established lunar simulants.

Abstract

Under low effective stress conditions, the often-neglected soil tensile strength may significantly influence the behavior of lunar and Martian regoliths. Therefore, regoliths’ tensile strength in extraterrestrial environments may pose many challenges for future space operations and exploration such as in situ resource utilization (ISRU). To explore regolith tensile strength for physical modeling on Earth, this study assessed the efficacy of the split box direct tension test method in measuring small amounts of tensile strength of F-75 Ottawa sand with added crystalline silica powder (CSP) at 30% and 50% by weight, and more commonly used lunar simulants JSC-1A and GRC-3. Findings indicate an increase in tensile strength with increasing regolith density. However, JSC-1A and GRC-3 forced open the split box prior to applied loading at low relative densities. This highlights the importance of considering system friction and lateral forces in the specimen during analysis. Although split box testing provided reasonable estimates of tensile strength, the testing method is more appropriate for soils with significantly greater tensile strengths than those tested here.

Practical Applications

NASA’s in situ resource utilization (ISRU) efforts deal with the challenges from the lack of immediate access to supplies during space exploration and habitation. To use lunar regolith as an in situ resource, a more robust understanding of the strength properties is required. Due to the low gravitation forces acting on the lunar surface (roughly $1/6$ the gravity of Earth), it is hypothesized that soil tensile strength may play a larger role in lunar regolith strength properties. Therefore, a simple method for testing the tensile strengths of lunar regolith is needed. To accomplish this goal, this research modified the simple split box test method to reduce the system friction by more than 10-fold for measuring the tensile strength of extraterrestrial regolith simulants. Due to the reduction in the system friction, low tensile strength soils that naturally fall into a conical shape forced the split box to open prior to loading. Therefore, this research indicates that prior split box testing may have reported higher than actual tensile strength values, and that future split box testing should consider the action of lateral earth pressures along the failure plane, especially in soils predicted to have low tensile strength.

The leading-edge flap (LEF) on a delta wing is an attractive tool because it can regulate the vortex strength and structure by changing the separation position along the leading edge. This paper aims to evaluate the control effectiveness of the LEF through numerical approaches. A moving mesh technique, which is developed based on the radial basis function method, is adopted to simulate the deflection process of the LEF. Static and dynamic deflections are studied numerically to explore their effects on the aerodynamic performance and the vortical structure of the delta wing. The results show that the asymmetric deflection of the LEF leads to an asymmetric distribution of the leeward vortices, which generates nonzero lateral forces and moments varying linearly with the deflection angle. During the dynamic deflection process, the flap deflecting downward fed more vorticity to the primary vortex on the same side, and vice versa. The amplitude of the lateral force and moment changes linearly with the deflection angle by which their frequencies are determined.

Compared to the fixed-geometry inlet, the supersonic variable inlet improves its wide Mach number operating capability and efficient cruise property, mainly by regulating the throat area. However, the magnitude of its performance improvement is constrained by the unstart and restart boundaries. To broaden the operating boundaries of the inlet, a distributed bleed system capable of suppressing the shock-dominated internal flow was developed. Its control effect and influence mechanism were investigated by an unsteady numerical simulation with a dynamic grid technique at freestream Mach numbers of 2.5, 3.0, 3.5, and 4.0. The results indicate that the proposed distributed bleed system can adaptively tune the bleed ratio to maintain the dynamic stability of the inlet’s flowfield according to the varying wave system in the internal duct. Additionally, the inlet’s unstart and restart boundaries were significantly improved. The unstart internal contraction ratio (defined as the area ratio of the duct entrance and throat, denoted as ICR) increased from 1.96 to 2.18, and the restart ICR increased from 1.37 to 1.47 when the freestream Mach number was 3.0. The mechanism of improving the unstart boundary is that the distributed bleed system alters the boundary-layer characteristics of the bleed region, making it thinner and fuller, with stronger resistance to reverse pressure gradients. The mechanism of improving the restart boundary is that the distributed bleed system reduces the spillage by increasing the inlet’s throat flow capacity, making the separation shock at the duct entrance impinge earlier on the cowl lip than that of the uncontrolled inlet. Moreover, the control effects of the distributed bleed system at different bleed ratios and under a wide Mach number range were investigated, which will provide guidance for practical engineering design and application.

A bird strike is a major risk to aeroengines. It not only causes the damage of blades and casing due to impact, but also leads to an outage of the engine due to the deposition of bird fragments. A finite-element method–discrete-element method (FEM-DEM) adaptive transfer method has been proposed and developed to simulate a bird strike by transferring FEM elements of the initial intact bird to DEM elements of bird fragments, which is superior to that of smoothed particle hydrodynamics (SPH). The cohesion of bird fragments provided by blood and fat could be simulated by the DEM algorithm. The characteristics of deposition, agglomeration, and motion track of bird fragments are simulated by three specific cases with different impact sites of the bird at two different velocities of the bird and rotation speeds of the engine. It reveals that deposition and distribution of the bird fragments are highly affected by the impact sites and velocities. The most serious deposition of the bird fragments in the inner channel occurs when the bird impact direction points to the stator blade cascade with lower impact velocity of the bird and rotation speeds of the engine. The bird fragments distributed more dispersed in this case. This simulation method provides an economic and efficient simulation technique for the design and damage assessment of the engine.

When space truss-membrane structures periodically enter and exit the sunshine area and the shadow area, it may generate thermally induced vibration in the interaction between thermal shock and structural deformations. Isogeometric analysis is applied to the thermal-structural coupling dynamics analysis of space truss-membrane structures. Established models of beam and membrane elements can properly describe the large deformation of the flexible structure under thermal shock. The thermal governing equations of thin-walled beams and membranes are derived, and isogeometric discretization is performed. Equations of motion and thermal governing equations are solved iteratively by the generalized-$\alpha$ scheme. The correctness of the model and the solution algorithm is verified by the thermally induced vibrations of a space cantilever thin-walled tube. Finally, a model of the truss-membrane structure is devised, and its dynamic response at different sun incident angles is analyzed. The results show that the maximum quasi-static deformation of a truss-membrane structure occurs when the sun shines vertically. Thermal flutter may occur when the space boom axis points away from the sun and the large deformation of the membrane under thermal shock causes lateral disturbance to the truss. The results indicate that the proposed isogeometric analysis can be used to evaluate the dynamic stability of space truss-membrane structures in orbit.

Abstract

For a pilot, the angle of attack (AoA) estimation may represent a possibility to cross-check the traditional cockpit indicators and also to obtain information whose benefits are widely recognized in many inflight situations. This paper proposes an AoA algorithm—named ACF—capable of detecting fast changes in the actual signal, that is a useful stall prevention capability, for example. The solution is based on the complementary filtering methodology and does not involve any airdata sensor output, with consequent saving of weight and cost reduction. In particular, it does not use the airspeed measurement, easily subject to failure and has been identified as the root cause of several fatal accidents. Consequently, the ACF shows the same (high) reliability as the used inertial sensors. Furthermore, the ACF algorithm requires neither information on the aircraft aerodynamic coefficients nor forecast wind information to be preloaded onboard, thus being time and cost effective. The ACF is also characterized by a low computational load, only needing a few dozen of algebraic operations. In this work, the ACF performances have been assessed using both a detailed simulation environment and real flight data. In the former case, an airliner following a sample maneuver is assumed, and a mean root mean square (RMS) less than 3° is guaranteed under all possible conditions of wind, gust, and turbulence. In the latter case, data collected in two different general aviation flights are used, providing a mean RMS of less than 0.3°.

Practical Applications

During piloting, the angle of attack (AoA) readings may represent a possibility to cross-check the traditional indicators and to obtain information whose benefits are recognized in many inflight situations. Unfortunately, the AoA sensors are exposed to several possible failures, often common to the other airdata such as airspeed and altitude. In addition, especially for small aerial vehicles, the systems cannot accommodate direct sensors due to size, weight, and power constraints. For these reasons, having an AoA sensorless estimation would be extremely important. A software solution (ACF) is proposed to estimate the AoA without using any external sensor, saving weight and costs, and with a strong improvement of reliability since the angle vanes—and especially the air probes—are generally identified as the root cause of several inflight accidents. Due to the fact that the ACF algorithm relies on inertial measurements only, it could even be implemented in a pilot’s personal device because it does not require any tuning or flight calibration. Furthermore, in spite of its simplicity, it can achieve good accuracy by detecting—in any weather and regardless of particular airplane or maneuver—fast changes in the actual AoA that may endanger the safety of the flight.

The design of nearly all Earth-orbiting spacecraft includes some sort of added shielding that protects the spacecraft against impacts by meteoroids and orbital debris. The effectiveness of the shielding is typically assessed using ballistic limit equations (BLEs) that are developed using data from high-speed impact tests on key spacecraft components and elements. These equations predict whether or not a particular system or structural element will sustain a critical failure following a specific impact event. As such, they are essential components of spacecraft system design as well as any quantitative spacecraft risk assessments that may need to be performed as part of that design process. Previous high-speed impact test programs have typically used medium-to-high density materials as surrogates for the kinds of materials that populate the orbital debris environment surrounding the Earth. However, with the advent of several new interplanetary spacecraft being designed, it is important to be able to predict and assess the performance of candidate spacecraft shields under impacts by projectiles made of materials that such spacecraft might be expected to encounter throughout their missions. To begin to address that issue, we assess how well a frequently used BLE, the New Nonoptimum (NNO) BLE, is able to predict the response of a commonly used shielding system when it is subjected to less dense materials that are often used as surrogates for icy meteoroids. In the end, it is shown that this BLE might require some modification when used to predict the response of such a system under the impact these kinds of projectiles.

In this study, streamwise vortex structures produced by wishbone vortex generators (VGs) and their interactions with a 30° backward-facing ramp (BFR) are investigated to understand their impact upon BFR flow separation behavior at $\mathrm{Re}=3\times {10}^6$. In particular, the effects of skewing the VGs relative to the free-stream are considered and results compared to those reported for vane-type VGs previously. Near-field formations of and interactions between the primary streamwise, horseshoe, and edge vortices are clarified along with the stagnation points and recirculating regions, where they are sensitive towards the VG skewness angle. Streamwise vortex-core strengths, trajectories, and other characteristics are also quantified. While the formations and behavior of streamwise vortices by wishbone and vane-type VGs imposed on a BFR are generally similar, there exist certain key differences. Comparisons with earlier studies show that vane-type VGs tend to produce far more persistent streamwise vortices with stronger circulation levels, while wishbone VGs tend to produce weaker streamwise vortices that convect slightly closer to the BFR surface. Interestingly, numerical predictions show that multiple streamwise vortices produced by a wishbone VG array interact more significantly than its vane-type counterpart even at a small skewness angle of 10°. This leads to much larger reattachment length reductions along the BFR span and potentially offers considerable pressure drag reduction levels.

An analytical nonlinear model for elastic planar motion of a tethered spacecraft (TS) is presented and used to investigate the mode characteristics and analyze the nonlinear vibrations induced by the orbital maneuver. Dynamic characteristics such as frequency densification caused by variation of solar panel length, mode localization, and mode interchanging phenomenon caused by acceleration variation are investigated. Two prerequisites and the mechanisms for the occurrence of the complete mode interchanging phenomenon are discussed in detail. Nonlinear phenomena were observed such as hardening, jumping, and third-order superharmonic resonance. Dynamic responses of TS with and without considering the geometrical nonlinearity of the tether were obtained and used to show the influence of the tether’s nonlinearity on the whole system. The influence of the tether’s nonlinearity was strengthened with the increase of the excitation amplitude. Numerical results showed that the geometrical nonlinearity of tether has a significant influence on the dynamical responses and residual vibrations of the whole tethered spacecraft.

Conventional approaches of design variable generation in nacelle design cases are usually proposed from either of two perspectives: (1) utilizing geometric resemblance to airfoil shape; or (2) selecting flexible geometric parametrization so that any potential geometry can be produced. However, the relationship between geometry and aerodynamic performance objectives is not emphasized in these approaches: design variables are likely not carrying the information that is most needed for aerodynamic optimization, as is the case in nacelle cowl design of drag reduction. In fact, an ideal aerodynamic information should be related with the move of samples toward better aerodynamic performance in design space. As an attempt to fix this issue, this paper proposes a new design variable generation method by extracting objective space (i.e., the aerodynamic performance of samples mapped from nacelle design space) information via linear discriminant analysis (LDA). By accomplishing this “objective space filtration,” the aerodynamic performance improvement potential is injected into design variable generation. Proof-of-concept studies of nacelle cowl design are provided in the paper demonstrating a more direct relationship that links: (1) LDA-aided design variables versus cowl drag, and (2) LDA-aided design variables versus shock strength and laminar length, respectively. LDA is used in the studies as a classic technique of classification, so that the analysis of objective space information finds a way to have impact on design space. The new approach outperforms a controlled test realized by Proper Orthogonal Decomposition in terms of reflecting aerodynamic information in design space, in that a localized improvement of aerodynamic objectives can be more clearly seen in the newly generated design variables. The result shows that the new design variable generation helps improve aerodynamic design and that it can be further extended to other cases beyond the scope of nacelle cowl design.

This paper investigates the prescribed-time tracking control problem for spacecraft with parametric uncertainties and deferred full-state constraints. Such kinds of constraints universally take effect some time after the system maneuvers, instead of being fulfilled initially, and are regularly encountered in practical applications. To achieve the deferred constraints, the modified states are incorporated into the control system with the aid of the state shifting function, thus removing the feasibility condition required by the traditional barrier Lyapunov functions. Considering that most of the existing results are only able to achieve tracking control in an asymptotic or finite-time manner, in this paper a prescribed-time control strategy for a spacecraft tracking control system is therefore studied. Synthesized with deferred constraints, the settling time can be roughly specified in a predetermined time bucket. Moreover, the parametric uncertainties can be addressed with a neural-based minimum-learning-parameter (MLP) algorithm. Finally, simulation results are provided to illustrate the efficacy of the designed control strategy.

In this work, the influence coefficient method together with a sliding plane method was used for an efficient analysis of fan acoustic flutter. The method requires the computational domain to be split to two zones: the rotating fan zone and the stationary intake zone. The subdomain interface is connected using the sliding plane method. The split of the domain is necessary for a stable solution. The time-accurate dual-time stepping method was applied for solving the blade vibration-induced unsteady flow. To demonstrate this method, a wide chord fan with two bell mouths of different duct lengths was chosen as the test case. The effects of acoustic reflection from the bell mouth on the fan aerodynamic damping were investigated by varying fan blade vibration frequency at 80% speed. The acoustic reflection has a noticeable influence on the fan aerodynamic damping only when the blade vibration frequency is close to the duct resonance frequency. The results also indicate that the slightly cut-on scattered modes impact the fan aerodynamic damping as well. The proposed method is validated by comparing with traveling wave method solutions.