Journal of Aerospace Engineering

A typical Hall thruster is powered by a constant-voltage power supply. In this study, to ensure the normal operation of Hall thrusters at lower power conditions, we applied pulse modulation technology to supply power to a 600-W Hall thruster. The results show that, when the input power of the thruster was within the rated power range, the thrust and efficiency of the thruster were slightly reduced under pulsating mode compared with those of direct current (DC) operation. However, the thrust-to-power ratio of the thruster under pulsating mode was superior to that under DC operation at low power ($<350\mathrm{W}$). The thrust-to-power ratio under pulsating mode was improved to $77.5\mathrm{mN}/\mathrm{kW}$ at 210 W compared with that under DC operation ($70\mathrm{mN}/\mathrm{kW}$). Meanwhile, we found that the stable operating boundary of the thruster under pulsating mode significantly widened at high frequencies. A 600-W Hall thruster under pulsating mode can operate at a lowest power of 66.8 W, which is an order of magnitude lower than the rated power. The thrust and anode efficiency under pulsating mode at 68.8 W were 4.3 mN and 7.8%, respectively.

Abstract

A spatial, discrete Michell truss to support a cantilever load is proposed in this paper. The node placement scheme is designed to iteratively generate the topology of the joints for the tensegrity structure. A recursive formula for the member forces is derived to compute the minimum volume of material needed to support a cantilever load. An example is provided with comparisons to the planar result special case. This example demonstrates that the benefit of supporting out of plane loads comes at a minor cost of additional mass. Further, a practical realization of the planar special case is compared to the proposed truss to demonstrate improved mass efficiency over conventional trusses.

Practical Applications

Tensegrity structures have been proven to be the optimal mass solution to the canonical loading conditions: bending, simply supported, torsion, compression, and tension. Further, the geometry of tensegrity structures can be efficiently manipulated through member length actuation. This allows for compact storage and subsequent deployment, which is economically attractive for space-based ventures where volume and mass are a precious commodity. The spatial discrete Michell truss in this work enjoys similar deployment capabilities, as the structure can wrap about its circular base when not in use. This work enables a practical realization of an optimal planar tensegrity structure, at the cost of marginal mass increase. This structure can support out of plane cantilever forces, which is not permissible in planar theory.

The effect of microporous surfaces on the secondary instabilities in a supersonic boundary layer was investigated using linear stability theory and nonlinear parabolized stability equations. The linear stability results indicate that the porous surface significantly enlarges the growth rate of the first mode disturbance and advances the transition. A two-dimensional disturbance is more sensitive to the porous surface than is an oblique disturbance. Nonlinear stability analysis suggested that the two classical secondary instabilities, namely the fundamental and subharmonic resonances, lead to transitions when the disturbance amplitude is high. The subharmonic resonance is stronger than the fundamental resonance. A porous surface enlarges the growth rates of the primary and secondary instabilities and promotes the transition. As the porosity coefficient or pore radius increases, the transition onset position gradually moves upstream. However, the fundamental resonance is more sensitive to the porous wall than is the subharmonic resonance, and the porous surface affects the secondary instabilities in different ways. For the fundamental resonance, the porous surface not only enlarges the amplitude of the primary mode, but also directly destabilizes the fundamental disturbance. These two effects cause the fundamental-type breakdown to occur much earlier. For the subharmonic resonance, the destabilization of the primary mode is dominant, and the subharmonic disturbance is not sensitive to the porous wall boundary conditions.

Compact tether-based actuation is a suitable approach for the deployment of femto/picosatellite bodies from CubeSats using ultrasmall electrodynamic tethers for fuel-free propulsion and deorbiting. Despite the advantages of tethered satellite systems, control technologies for these have yet to mature in several domains including robustness to structural faults and unmodeled dynamics. A proposed solution for the identification of disturbances to tethered satellite dynamics is to use a data-driven algorithm to learn the system’s behavior over previous orbits and then provide an estimated prediction for the evolution of system states. To achieve the goal of state prediction via a globally linearized system model, this paper employs the Koopman operator constructed from observed dynamics to extrapolate future motion of a tethered subsatellite subject to unknown disturbances while being deployed from its mothership. Numerical simulations of the constructed model versus the nonlinear model of the tethered satellite system demonstrate the effective prediction capabilities of the proposed Koopman operator-based numerical algorithm for the general flight characteristics many orbits into the future.

Wind tunnel testing and accompanying numerical analyses are the essential methods to improve and optimize the aerodynamic design of helicopters. The focus is on improving the aerodynamic performance and the handling qualities. Of special interest is the parasitic drag as a major factor for the power consumption and the wake flow caused by the rotor head, which can interact with the helicopter tail and influence stability. For this purpose, design modifications and their influence on the aerodynamic characteristics of the helicopter as well as on the flow field are investigated. The design modifications are evaluated by force and moment measurements, stereo particle image velocimetry, and high fidelity computational analysis. This allows us to compare the different investigation methods and design configurations to each other. The investigations are performed on the Kopter AW09, a 2.8-t-class single engine utility helicopter. Design modifications of the upper fuselage are considered for the presented investigations.

This paper investigates the load-carrying capacity and failure mode of carbon fiber/epoxy resin T700/QY9611 composite T-joint with initial debonding damage under tensile load. The numerical results were verified by experiments. Based on the verified finite-element model, the effects of the length and location of initial debonding damage on the properties of the T-joint were further studied. It was found that the carrying capacity is closely related to the failure mode. If the initial debonding damage length is in the range of 0–5 mm, crack propagation does not involve the filling area, which is a tensile failure with a high failure load of about 20 kN. For the specimens with damage length greater than 5 mm, the shear stress causes obvious damage within the filling area, and the crack extends to the soleplate/L-rib interface, which is a mixed failure of tension and shear and the failure load decreases by 47.5%, to only around 11 kN. However, the load-carrying capacity of the T-joint is less affected by different debonding damage lengths under the same failure mode. The research results have important reference values for damage assessment and repair tolerance of composite T-joints.

This study investigated the effects of powered engines on the stability boundaries of a box-wing configuration. The governing equations were obtained using Hamilton’s principle. The equivalent effects of engine’s mass and thrust were simulated using a concentrated mass which was subjected to a follower force. The structural dynamics of the box wing were modeled using two cantilever beams which were attached to each other through two longitudinal and torsional springs at their tips. The aerodynamic loads applied on the wings were calculated using Wagner’s unsteady aerodynamic model. The time-dependent and the parameter-dependent integral parts were eliminated using the multistep semianalytical method. To validate the developed aeroelastic model, the flutter speed and frequency of conventional wings and box wings were obtained, and very good agreement was observed. The analysis showed that the stability of the wing can be affected when the effects of engines are considered. Furthermore, the engine placement and thrust effects on the box-wing flutter speed and frequency were considered.

In this paper, static and dynamic characteristics of a labyrinth seal were studied by fluid–solid–thermal coupling analysis (FSTCA). Deformation and stress distribution of nonmetallic labyrinth seals (i.e., PEEK and PEEK-CA30) were further analyzed to investigate the effects on leakage and rotor stability compared to that of metallic material (i.e., AL alloy). The results show that a higher pressure drop of airflow at the last tooth along the airflow direction suggested a more remarkable energy conversion to kinetic energy. Airflow velocity at the tooth root was obviously lower than that at the tooth tip because of vortex in tooth cavity. Maximum deformation of the seal occurred at tip of the first tooth, while equivalent stress mainly distributed at the root. Both axial and radial deformations led to deformations of seal cavities and upward movements of the teeth, resulting in greater radial clearances. PEEK-CA30 had the lower radial and axial maximum deformations compared to PEEK and AL alloy, which in turns caused a lower leakage with increasing inlet temperature and inlet/outlet pressure ratio and a higher effective damping coefficient under different vortex frequencies, indicating a better seal performance and a better stability of rotor system. The equivalent stress for AL alloy deeply depended on inlet temperature of airflow due to higher elasticity modulus, higher thermal conductivity and higher thermal expansion coefficient, which made the seal more likely to fail under higher inlet temperature.

In the flight process of a rudder wing structure in a hypersonic environment, a large temperature gradient is generated due to severe aerodynamic heating, leading to structural strength failure or ablation failure under the coupling effect of thermal stress aerodynamic force. Therefore, this paper proposes a method to study the ablation behavior of the hypersonic rudder wing structure. Based on the basic theories of hypersonic aerodynamics and structural mechanics, a fluid–solid–thermal coupling model of the rudder wing structure in a hypersonic environment is established. The temperature distribution and deformation of the rudder wing structure in the service state are obtained. In addition, the finite element analysis method based on automatic remeshing simulates the ablation of the rudder wing structure in a hypersonic environment. Under inflow conditions, the leading edge of the rudder has high temperatures and large deformation. The ablation simulation results are compared with the wind tunnel test results, and the results show that the proposed method can be applied to the study of service behavior in a hypersonic environment.

The parafoil has excellent gliding ability and operability for recovery of the first-stage booster, which makes the landing point of the first-stage booster controllable. However, during the recovery process, the wind field changes in real time due to changes in altitude and turbulent wind, which has a great influence on the trajectory and landing point of the parafoil recovery system. In order to realize accurate homing of the parafoil first-stage booster combination (PFC) in a complex wind field, an online trajectory planning algorithm based on the quartic Bézier curve is proposed in this paper. This algorithm combines the Bézier curve with the multiphase homing method and proposes a method to calculate the control points of the Bézier curve according to the time-varying wind field. First, a six-degrees-of-freedom model of the PFC is established, and the motion characteristics of the PFC are analyzed. Then, the drift of the PFC caused by the wind field is calculated, and it is superimposed on the target landing point to obtain the virtual landing point. On this basis, the quartic Bézier curve is used to plan the homing trajectory. The simulation results show that the algorithm has high computational efficiency for online trajectory planning. It can realize accurate homing of the PFC in complex wind fields with continuous unilateral brake deflections, which makes it conducive to practical application.

For a hypersonic vehicle powered by modular scramjet, integration of the airframe with the engine is critical to generating positive net thrust. In this study, a design method that combines the ideas of streamline tracing and trimming is generalized to design the modular shape-transition nozzle for integrating with the axisymmetric vehicle after-body. As an example, the design of a modular shape-transition nozzle with three modules for an axisymmetric vehicle is presented, and the computational fluid dynamics (CFD) approach is adopted to obtain the detailed flowfield and performance of the modular nozzle. Smooth expansion of the burner-exits gas along the flow direction is involved in the modular nozzle. The flow interactions between any two module exhausts, as well as the flow interaction of the modular jet with the external freestream, are all appearing at the trailing edge of the modular nozzle. Moreover, the modular nozzle produces a net thrust coefficient of 0.8536 and an integrated thrust coefficient of 0.7896 for the vehicle. The mixing of the modular nozzle jet with the freestream is enhanced with the trimming of the separate nozzle module, an approach which can reduce the infrared radiation signal.

Despite the increasing interest in unmanned aerial vehicles (UAVs), their adoption in commercial flight operations invariably meets with skepticism, mainly on the basis of safety and reliability concerns, as well as poor payload and endurance of available—generally rotary-wing—platforms. However, specific missions exist where higher-weight UAVs may be employed, specifically to serve wild or disadvantaged areas, far from crowded regions, and transporting medical aids or food. Clearly a niche too small to be considered profitable as a market for a new design by industry, this requirement can be fulfilled through the partial redesign of an aircraft in the light sport aircraft (LSA) weight class. Based on a set of specifications discussed with Médecins Sans Frontières (MSF), the present paper analyzes the feasibility of a mission where medical aids are carried over a prescribed route by means of a UAV, and parachute-dropped on the target area. A candidate for the proposed mission is found in an existing LSA. Its optimal use and the corresponding retrofitting steps to fit within the prescriptions of the mission are proposed and critically discussed.

The reliability during aeroengine operation is significant in monitoring its time-varying operational status and providing valuable reference for both operation and design. To study aeroengine time-series operational reliability and parameter feature effects in whole flight stages, a tree model is developed to discuss the operational reliability of aeroengine through residual thrust and feature importance analysis. 24 parameter features from an aeroengine (CFM56) operation are applied to investigate the required thrust and residual thrust of aeroengine through quick access recorder (QAR) flight data. Characterizing aeroengine load case by the cosine similarity of aeroengine feature sets and operational reliability of aeroengine are conducted by combining cosine similarity and residual thrust of aeroengine. Further, XGBoost (extreme gradient boosting) model coupled with the Shapley additive ex-planations (SHAP) algorithm is constructed to enhance model interpretability analysis, in which the importance and influence direction of aeroengine features are analyzed. In this study, features similarity and importance were analyzed using aeroengine QAR data from a single flight. The results showed that (1) the operational reliability of aeroengine greatly varies under different flight stages, (2) the load case of aeroengine undulate greatly under take-off and descent stages, but is stable in the cruise stage, and (3) the feature importance and influence direction are indifferent among residual thrust and three aeroengine parameter sets, showing the complexity of aeroengine operation. Based on the operational and applicable restriction data, the efforts of this study provide a new analysis method for aeroengine to improve the design and operation by analyzing operational reliability.

Drag-free control technology plays a key role in space science tasks such as microgravity science, space basic physics verification, Earth observation, and spacecraft high-precision navigation. At present, a drag-free controller is designed in the time domain aimed at the full frequency band by many scholars, which cannot meet some of the high-precision task requirements under specific frequency constraints. Therefore, some scholars use robust control and quantitative feedback control to design a controller in the frequency domain indirectly, but the design process is complicated. Moreover, the satellite tracking test mass strategy is generally used, and it limits the control capability of a drag-free satellite with double test masses. For these problems, a design method of a finite frequency domain controller based on Generalized Kalman-Yakubovich-Popov (GKYP) lemma combined with frequency separation control strategy is presented in this paper, which covers the dynamic model, design of the finite frequency ${\mathrm{H}}_{\infty }$ controller, proof of stability, and performance analysis. Finally, the numerical simulation analysis is carried out. The simulation results show that the controllers are effective.

In this effort, experiments are conducted for a small-scale remotely controlled (RC) helicopter in both hovering and forward flight modes. An open wind tunnel type is used as a source of forward airflow for the helicopter model with flow speed ($v_{\infty }$) ranges from 3 to $12\mathrm{m}/\mathrm{s}$. A test rig was designed and manufactured to provide safe and reliable averaged thrust measurement in hovering and forward flight regimes. The helicopter is controlled via an adhoc-designed graphical user interface program using a microcontroller along with sensors to measure the blade collective pitch (${\theta }_0$), rotational speed ($\mathrm{\Omega }$), and shaft angle of attack (${\alpha }_s$). The averaged thrust for the helicopter is measured and compared with blade element theory and an unsteady model based on strip theory. Measurements are conducted for six rotational speeds and five collective pitch angles in hovering mode. Additionally, forward flight measurements were conducted for five rotational speeds, five collective pitch angles, two forward airflow speeds, and two shaft angles of attack. Uncertainty analysis is performed to certify a statement of the obtained results by combining both bias uncertainty from measuring instrument specifications and precision uncertainty through experimental repeatability. Results can be used for preliminary sizing of rotors for such small helicopters in both flight modes.

The effects of a spike on the aerodynamic interference reduction for a multibody configuration were investigated numerically. A two-stage-to-orbit (TSTO) vehicle was selected as a classic multibody configuration. A spike was placed on the nose of the orbiter to change the position and shape of the shock wave. The peak pressure and peak heat flux on the booster upper surface were reduced by 39.5% and 21.5%, respectively, after the spike was installed. The peak pressure and peak heat flux on the orbiter nose were reduced by 78.2% and 28.4%, respectively. The total drag coefficient of the whole model was reduced by 27.3%. The results show that a spike is able to reduce the shock–shock interference and the shock wave–boundary layer interference (SWBLI) between two stages without changing the aerodynamic configuration of the aircraft.