Journal of Aerospace Engineering

In this paper, generating symmetric periodic orbits around a binary asteroid and optimal maintenance of spacecraft on these periodic orbits are investigated. Due to the irregular shape of asteroids, numerical methods are needed to generate periodic orbits. In this regard, the 32 families of symmetric periodic orbits were calculated by grid search and shooting methods utilizing the restricted three-body problem as a basic model. The irregularity of the shape of asteroids and solar radiation pressure are considered perturbations. These perturbations should be compensated by a thruster with a continuous thrust through the continuous control strategy, i.e., the linear quadratic regulator (LQR) method. The effect of different matrix coefficients in the performance of the LQR method is examined. The results show that the LQR method could appropriate control spacecraft on periodic orbits with specular shapes. Also, the practical feasibility of a continuous thrust control strategy is discussed.

In this paper, we present an adaptive control strategy for the output-constrained attitude tracking of rigid spacecraft under unknown inertia and disturbances. The designed controller is recursively constructed under the framework of the backstepping method, which involves an auxiliary stabilizing law and a practical control law. In the auxiliary stabilizing law design, the log-type barrier Lyapunov function (BLF) is embedded to tackle the time-varying output constraints. In addition, in the practical control law design, the element-wise parametric adaptive laws are adopted to identify the unknown inertia and disturbances by introducing a linear mapping operator. The stability evaluation indicates that the overall closed-loop system is semi-globally uniformly ultimately bounded (SGUUB) and all error variables can eventually regulate to the minor fields about zero. Meanwhile, the spacecraft attitude tracking errors can be preserved in the preassigned output constraints. Lastly, the efficiency and strong robustness of the presented control strategy are demonstrated through competitive simulations.

A new generation of transport aircraft increasingly adopts T-tail configurations for their excellent aerodynamic, operating, and structural performance. However, flutter of T-tail configurations is a serious dynamic instability problem caused by aerodynamic and structural interactions between the vertical stabilizer and the horizontal stabilizer. The horizontal tailplane (HTP) is mounted on the vertical tailplane (VTP). Bending and twisting of the VTP induce rolling, yawing, and in-plane motion of the HTP, which has a significant effect on T-tail flutter. In this paper, a multiple input/multiple output (MIMO) linear quadratic Gaussian (LQG) controller is designed to suppress flutter for a T-tail. The state-space equations of the aeroservoelastic (ASE) T-tail model, which are composed of the state-space equations of the aeroelastic system and the actuator dynamics, are derived for the design of the control law. The controller is synthesized by posing and solving a weighted optimization problem with the goal of making the rudders collaborate deflection for flutter mode suppression and expand the boundary of the T-tail flutter. A thorough analysis of the performance achieved by the closed-loop system was performed through numerical simulations. The numerical results demonstrate that the proposed LQG controller can effectively suppress the T-tail flutter and expand the flutter boundary.

This study abstracted the prediction of the aerodynamic coefficients of an iced airfoil as a mapping from the iced airfoil space to the aerodynamic coefficient space. Thus, a deep network called Airfoils2AeroNet that can learn this mapping was established based on Deep Operator Network (DeepONet). The deep network consists of a branch network for encoding the iced airfoil images and a trunk network that learns a nonlinear mapping from the one-dimensional aerodynamic coefficient function input to $p$-dimensional outputs. The branch network consists of deep convolutional neural networks (CNNs), and the trunk network consists of fully connected neural networks (FNNs). Then the network was trained and tested on iced airfoils based on NACA 0012 airfoils. Comparing the prediction results of Airfoils2AeroNet with those of the conventional direct CNN network, the network proposed in this paper has a strong advantage for generalization. Unlike the traditional CNN, which can only predict the aerodynamic coefficients at fixed flow conditions consistent with the training data, the network can flexibly predict the aerodynamic coefficients at different flow conditions. Finally, the influence of the structure of the branch network and trunk network on the prediction results was analyzed.

This paper addresses the fixed-time, six-degrees-of-freedom control problem for the distributed spacecraft formation system subject to actuator saturation, considering limited communication resources. Since only a fraction of the follower spacecraft can contact the virtual leader, a distributed fixed-time observer is designed to reconstruct the leader’s information. Subsequently, a distributed fixed-time, six-degrees-of-freedom controller is established, compensating the actuator’s magnitude and the rate saturation problem with an auxiliary system. In particular, this paper proposes a novel event-triggered mechanism in the framework of the “adding a power integrator” technique, avoiding the unnecessary occupancy of communication resources significantly. Finally, numerical simulations demonstrate the advantages of the theoretical results.

The shortcomings of the concurrent subspace collaborative optimization (CSCO) method for structural layout synthetical optimization problem have been analyzed in the current paper. Based on the complex method principle and the special form of structural layout synthetical optimization problem, a new system-level coordination strategy has been proposed to overcome these defects and improve the CSCO method. The computational stability of the CSCO method has been improved by this strategy. For the sake of distinction, the improved method is called the concurrent subspace complex coordination (CSCC) optimization method. Two examples, including a 25-bar truss structure and a composite stiffened panel structure, have been conducted to demonstrate the effectiveness and reliability of the CSCC optimization method.

Nacelle cowl design is dedicated to generating a three-dimensional surface with optimal drag performance. Most previous approaches focused primarily on the design method driven by streamwise profiles, which introduced no additional aerodynamically significant variables into the geometric constraints. Therefore, to explore more design spaces for drag reduction of the nacelle cowl, this paper proposes a design refinement method by adding a circular curve that passes through the locations of maximum thickness of given streamwise profiles at various angles toward the nacelle axis, which are regarded as the design variables. A proof-of-concept study was modeled using a combination of computer aided design, Latin hypercube sampling, computational fluid dynamics (CFD) simulation using Reynolds-averaged Navier–Stokes equations closed by a $k\text{-}\omega$ shear stress transport model, and a kriging surrogate model, in order to search for the optimal solution in design space. The result of optimization shows that the aerodynamic performance of the nacelle cowl was optimized significantly, with a 28.50% drag coefficient reduction with only three angular variables taken to create the design space, illustrating the feasibility of the proposed approach.

A sliding mode (SM) re-entry control system is proposed for an assured crew re-entry vehicle (ACRV). The vehicle has a low lift/drag ratio (0.3). The trajectory controller drives the ACRV in a predefined track. The reference vectors for the attitude controller are generated by the trajectory controller and navigation system. The feedback SM controller is designed to accommodate the structural parametric uncertainties and environmental disturbances. The SM controller is optimized by both the genetic algorithm (GA) and the simulated annealing (SA) methods. A comparative study is done between both optimization methods and the efficiency of the SA method is proved. The system is subjected to a sensitivity study in terms of various thrust torque profiles. A robustness check is done by changing the system parameters. A pulse width pulse frequency (PWPF) modulator modulates the output torque dictated by the attitude controller. The simulation results demonstrate the proposed method’s effectiveness in achieving the various required parameters.

The thermal environment behind the integrated strut of the afterburner is very harsh. A double-wall structure is set at the integrated strut flame stabilizer for cooling. The multiobjective optimization of the double-wall cooling structure is carried out by using the RBFNN-coupled nondominated sorting genetic algorithm II (NSGA-II), and the complex interaction between the film outflow and the mainstream was explored based on the numerical simulation results. The effects of geometric parameters and external working conditions on the cooling performance of the double wall are studied based on the Sobol method. The results show that the upward spiral vortex behind the strut will press the film outflow to form an irregular air film on the wall, effectively protecting the trailing edge of the strut. The optimized double-wall structure has a better air film coverage effect, and the comprehensive cooling performance has been improved to a certain extent. Compared with the geometric parameters, the mass-flow rate of the cooling gas per unit area has a greater impact on the comprehensive cooling performance of the double wall. The temperature ratio of the main and secondary flows has little effect on the comprehensive cooling performance of the double wall.

An experimental investigation of the aerodynamics and near-field tip-vortex flow field behind tapered backward- and forward-swept wings with a stationary ground effect was conducted at Reynolds number $(\mathrm{Re})=1.81\times {10}^5$. The results showed a large lift increase of 26.7% and 12.3% for the backward-swept wing (BSW) and forward-swept wing (FSW), respectively, with reduced ground clearance, along with a significant drag reduction of 45% and 30% for the BSW and FSW. For the BSW, a multiple-vortex system appeared in close ground proximity, consisting of a tip vortex, a corotating ground vortex, and a counterrotating secondary vortex. The ground vortex strengthens the tip vortex, whereas the secondary vortex negates its vorticity. For the FSW, the multiple-vortex system was not readily identifiable due to its unique geometry, which always keeps the inboard region of the wing at a close ground effect while leaving the tip region less affected by the ground effect. The root stall of the FSW also produced a continuously strengthening tip vortex with the increasing angle of attack. In contrast, the tip stall of the BSW led to a monotonically increasing vortex strength only up to the static-stall angle. Regardless of the wing model, the weak tip vortex also translates into a small lift-induced drag compared with the total drag. Finally, the lift force computed through the integration of the spanwise circulation distribution, inferred from the cross-flow measurements, at selected ground distances was also found to be in good agreement with the direct wind-tunnel force-balance data, with 101% and 98% consistency for BSW and FSW, respectively. The aerodynamics and tip-vortex measurements of both wing models outside the ground effect were also acquired to serve as a comparison.

In order to enhance the overall performance and operational efficiency of the integrated coaxial compound helicopter and turboshaft engine system, a preliminary study on the optimal rotational speeds was implemented based on multidisciplinary integration. First, the flight mechanics model of coaxial compound helicopter and the aerothermodynamic model of turboshaft engine were developed. Then, the minimum fuel flow of the turboshaft engine was selected as the optimization objective, and an integrated optimization method of optimal rotational speeds based on the golden section method was proposed according to fixed ratio transmission (FRT) and continuously variable transmission (CVT). Finally, the effect of the optimal rotational speeds on the overall performance of the integrated system was investigated, and the variation laws of the optimal rotational speeds under different operating modes were revealed. The results indicated that the optimal rotational speeds can decrease engine fuel flow by a maximum of more than 10%, which is conducive to promoting operational economy and overall performance. Moreover, the flight conditions, helicopter weight, and engine degradation significantly affect the optimal main rotor speed.

This paper presents wing design optimization with functional constraints using the proposed method. The proposed method is an equilibrium optimizer (EO), and it is motivated by controlling the volume mass balance method to calculate both equilibrium and dynamic states. A well-defined generation rate term is verified to strengthen the capability of the EO at exploitation and exploration. Here, the EO’s updating behavior is enhanced by Levy flight. Hence, it is named as Levy flight–based equilibrium optimizer (LFEO). The analysis device of aerodynamics is intended by combining the numerical nonlinear lifting-line method to attain the coefficients of lift and drag of the baseline wing. The proposed method is utilized for solving restricted optimization problems, and three various optimization problems are solved. The functional restrictions are connected to wing weight, wing plan form area, and root bending moment, which is included in the first optimization issue, and then the second optimization issue is created. The third optimization issue is attained by including to the second issue the state of level flight and minimum speed of the baseline unmanned air vehicle (UAV) at the maximum speed of level flight and obtainable power restrictions. It shows that describing the moment of root bending, wing area, and obtainable power restrictions at aerodynamic optimization issues may lead to various realistic wing-plane and airfoil shapes. Finally, the proposed method is implemented using MATLAB or Simulink and it is compared with existing techniques such as EO and genetic algorithm (GA).

Abstract

The distorted inlet flow that is a widespread flow phenomenon in compressors not only reduces the compressor performance but more seriously deteriorates the compressor stability, thus threatening flight safety. However, quantitatively determining the stability boundary is challenging due to the complexity of the distorted inlet flow. The effective static-pressure-rise coefficient in the stability criterion from Koch, a stability analysis method under clean inlet flow, was applied to explore the compressor stability under circumferential total pressure distortion. Using a series of single-stage axial compressor experiments for aspect ratio and solidity, this study investigated the influence of the distortion index on compressor stability. Subsequently, based on the analysis method under clean inlet flow, a new stability criterion under circumferential total pressure distortion was established to determine quantitatively the stability boundary. In addition, the stability criterion for determining the stability boundary was combined with the body force model for calculating the compressor performance to simulate a five-stage axial compressor under 120° total pressure distortion, and the results were in agreement with the experimental data to verify the stability criterion. Numerical simulation was used to show the internal flow field in the compressor, analyze the load at each stage, and identify the instability stage that induces stall, which can provide a reasonable solution for determining the stability boundary of a multistage axial compressor under circumferential total pressure distortion. The combination can serve as a convenient preliminary design tool to evaluate compressor stability using fewer computational resources.

Practical Applications

Most stability investigations adopt simple methods such as numerical divergence and specified parameters under circumferential total pressure distortion, which makes the stability boundary inaccurate. Since the stability analysis methods under clean inlet flow have been studied more deeply so far compared to that under distorted inlet flow, a new stability analysis method under circumferential total pressure distortion is established to quantitatively determine the stability boundary based on the stability analysis method under clean inlet flow. Therefore, in the preliminary design phase, the stability analysis method is combined with the model for calculating the compressor performance to scientifically evaluate the stability of the multi-stage axial compressor under circumferential total pressure distortion. Also, the combination simply requires fewer computational resources than the millions of grids in commercial software.

This study proposes a combined fault diagnosis scheme based on a recurrent neural network (RNN) and an $H_{-}$ observer for satellite attitude control systems (ACSs) in the presence of model uncertainties, external disturbances, and measurement noise. The ACS is decoupled into multiple independent channels, such that the residuals generated by the observers respond to the corresponding faults. A novel multilayer adaptive Gaussian recurrent neural network (MAGRNN) structure is developed as an approximator to estimate the lumped disturbance; a robust term is introduced to improve accuracy. Considering the actuator fault in the finite frequency domain and the fault-sensitive index, a set of $H_{-}$ unknown input observers (UIOs) is designed using the output of the MAGRNN-based approximator as compensation. The existence conditions of the approximator and observer are proposed and proved. The fault diagnosis results for three cases verify that the proposed method can be used for small fault detection and isolation.

Research into gust interactions has shown significant potential for extracting energy from the atmosphere. Notably, improvements to gust-energy extractions have been shown to be possible through aeroelastic tailoring. The present work uses a design-space exploration to quantify specific structural parameters that provide the greatest energy gains from discrete and continuous gust fields. Using a genetic algorithm and full-factorial parametric sweep, wing configurations that are stiff in bending but flexible in torsion were identified to provide the best energy gains for sinusoidal and 1-cosine gusts. These designs were also subjected to a continuous von Kármán gust field, showing the same trends as the discrete profiles. The influence of the elastic and mass axis locations was shown to be much weaker on the energy gain than the stiffness parameters. Overall, energy gains relative to the steady-gliding flight performance of the aircraft of 15% were shown to be achievable through aeroelastic tailoring.

In space applications, the pursuit of deployable structures with high packaging efficiency and low mass is essential. Tensegrity structures, prestressed networks of struts and cables, offer a promising approach. In this work, a Class-2 cylindrical triplex tensegrity bay is explored for use in booms, where select cables are allowed to change length in order to modify the overall shape of the structure. A novel approach to characterize the range of equilibrated shapes a triplex can achieve is first presented. The results are then used in combination with consideration of bending stiffness, packaging efficiency, and mass to explore the effects of different configuration parameters, namely the ratio of bay height to circumscribing radius ($h/r$), the member cross-sectional radii, and the number of bays. This design analysis offers new insight into how triplex tensegrity structures can be modified to achieve specific design objectives. Increasing $h/r$ for a single bay increases the range of achievable active tip motion as well as packaging efficiency. Increasing member radii decreases the range of achievable shapes and packaging efficiency while increasing stiffness and mass. When increasing the number of bays for a fixed overall height, a larger number of bays generally increases the range of achievable shapes and mass while reducing stiffness. Packaging efficiency was found to be discontinuous dependent upon the exact number of bays, but generally decreases with increasing number of bays. An example design study is conducted using the Canadarm2 for comparison and illustrates the potential utility of tensegrities for boom structures.

Compared with the traditional docking mechanism, a low-impact docking mechanism (LIDM) has a big advantage, that is, LIDMs can realize low-impact docking by controlling the movement of the driving rods in real-time. In this process, the load sensing ring (LSR) of the chaser spacecraft is adjusted in real-time according to the position and attitude of the target spacecraft, which requires that the LIDM has high transmission efficiency, so that the position and attitude of the LSR can be changed rapidly under the motion of the driving rods and respond to the position and attitude requirements of the docking control system to the LSR as soon as possible. In this paper, a transmission efficiency solution and an optimization method suitable for LIDMs are proposed. The LIDM dynamic model and transmission efficiency model are established, and an improved genetic algorithm with better convergence effect is proposed and used to optimize the transfer efficiency. The results of numerical examples in this paper show that the improved genetic algorithm has stronger global optimization ability when calculating high-complexity optimization problems. In the optimized LIDM configuration parameters, the size of LSR and the bottom reference circle are close to the lower limit and upper limit of the given size limit, respectively, and the instantaneous transmission force ratio is increased from 0.6477 to 0.9160. The reliability of the optimized results is verified in dynamic simulation software. This study provides ideas and methods for the design and optimization of subsequent high transmission efficiency mechanisms.

This paper proposes a robust finite-time attitude controller for spacecraft with input saturation by utilizing the full-order sliding mode control (FOSMC). Using the hyperbolic tangent function and homogeneity theory, a new full-order sliding mode (FOSM) surface was developed which makes the states of the system reach the equilibrium point in finite time. Based on the strengths of the designed FOSM surface, a robust attitude controller that makes the errors of the system converge to zero in finite time was developed for spacecraft without chattering and singularity problems. The effectiveness of the presented controller was verified via rigorous theoretical analysis and simulation results.

Adaptive chirp mode decomposition (ACMD) is a novel decomposition algorithm, which can decompose the multicomponent signals and identify its instantaneous frequencies (IFs) simultaneously. It is crucial for ACMD to predefine the decomposition parameter (i.e., the initial IF). However, the nonstationary signals with closely spaced frequency modes or heavy noise render it difficult to preset the initial IF based on the Fourier spectrum. To overcome the difficulty of predefining the initial IF of the ACMD, a tractable version of the ACMD, termed as the enhanced adaptive chirp mode decomposition (EACMD), is proposed in this study. The EACMD adopts an adaptive method based on the maximum entropy power spectrum to automatically determine the initial IF. A simulated dynamic signal is used to evaluate the characteristics and performance of the EACMD. Furthermore, the proposed EACMD is employed to identify the IFs of time-varying structures. The effectiveness of EACMD is illustrated and verified by using a 3-story shear building model and a cable with time-varying tension forces. The results show that the EACMD can overcome the difficulty of predefining the initial IF of the ACMD and is suitable for identifying the IFs of time-varying structures, providing more accurate results than the existing techniques.

The present study numerically investigates the aerodynamic performance of the morphing airfoil and wing with leading edge (LE) and trailing edge (TE) parabolic flaps. Reynolds-averaged Navier-Stokes (RANS) equations coupled with the Spalart-Allmaras (S-A) turbulence model are employed to simulate the flow fields. Firstly, the parameterization method for the morphing airfoil with LE and TE parabolic flaps is proposed. Secondly, the influence of several design parameters, namely, LE droop angle, LE droop position, TE deflection angle, and TE deflection position, on the airfoil aerodynamics are further explored. Moreover, the aspect ratio (AR) effect for the wing based on the morphing airfoil is also investigated. Simulation results demonstrate that the parabolic flaps outperform the articulated flaps in lift generation, drag reduction, and aerodynamic efficiency enhancement, with an improvement in $C_L$ and $L/D$ by 39.2% and 108.4% at $\alpha =4°$, respectively. A larger LE droop angle can increase the stall margin, whereas the LE droop position barely influences airfoil aerodynamics at small angles of attack. The increment of the TE deflection angle significantly augments the lift. As the TE deflection position moves downstream, the suction peaks at the LE and TE are decreased and flow separation is also delayed. Furthermore, a larger AR can achieve higher efficiency, whereas a smaller AR is more beneficial for flow separation suppression.

The identification of two-dimension distributed dynamic loads using partial structural responses is still challenging. In this paper, a method is proposed to identify the gradually varying spatial distribution functions and the time histories of time-space decoupled two-dimension distributed dynamic loads using partial structural responses. First, due to the properties of good decorrelation and energy compression by discrete cosine transform (DCT), two-dimension DCT is used to expand the spatial distribution functions with fewer coefficients than those by other transform such as discrete wavelet transform (DWT) in previous identification approaches. Under given DCT coefficients, the load time histories and structural responses are identified by the Kalman filter with unknown input (KF-UI). Then, the DCT coefficients are evaluated based on minimizing the errors between the estimated and measured structural responses. Finally, the spatial distribution function is reconstructed by the inverse two-dimension DCT with the estimated optimal DCT coefficients, and the load time history is identified by KF-UI accordingly. The novelty of the proposed method is that fewer DCT coefficients are involved to expand the gradually varying two spatial dimensions, which simplifies the difficulty of estimating many coefficients using partial structural responses. Numerical examples of identifying different dynamic loads verify the effectiveness of the proposed method.

For the first time, a two-node beam element with nine degrees of freedom (DOFs) was developed in this paper using a combination of modified couple stress and $k$th-order shear deformation beam theory to analyze the free vibration and static buckling responses of rotating bidirectional functionally graded porous (BDFGP) microbeams resting on variable elastic foundations, in which the whole structure was exposed to a magnetic field in the hygrothermal condition. The mechanical properties and the length-scale parameters varied in both the thickness and length directions, which in the novel porosity model and power-law indexes of the material were considered to be a function of the $x$- and $z$-coordinates. The present method’s accuracy was determined by comparing it with published findings. Additionally, extensive research was conducted to evaluate the effect of various parameters on the mechanical responses of the rotating BDFGP microbeam. The numerical results showed that porosity coefficient, elastic foundation parameters, and hygrothermal environment all significantly affect the rotating BDFGP microbeam’s free vibration and static buckling behavior.

This paper investigates an artificial gravity strategy using the spinning tethered spacecraft (STS), which is spun along orbital direction. An appointed-time prescribed performance (ATPP) attitude controller of tethered spacecraft with unknown external disturbance is proposed by backstepping technology. First, the compound thrust control scheme is presented to provide an ideal artificial gravity overload by the centrifugal force on the tether. Then, the novel performance function is devised to ensure the dynamic response of the system but also achieves appointed-time stability. Compared with the traditional disturbance rejection method, an asymmetric barrier function is introduced to suppress unknown disturbances without estimated information. Based on the spin control scheme and the proposed attitude control strategy, the system is proved to be globally uniformly appointed-time stable in the presence of unknown disturbance. Furthermore, applications to the STS are employed to show the effectiveness of the ATPP control approach for artificial gravity.

This paper mainly investigates the aerodynamic effect of two different average sizes and three specific nonaxisymmetric rotor tip clearance arrangements on the overall performance and loss mechanism of a four-stage low-speed research compressor (LSRC) at Shanghai Jiao Tong University. The stage was modeled for the steady simulations as a seventh entire annulus computation domain of the inlet vane and the first stage rotor row. The numerical method was validated using experimental data on the radial profiles of total pressure and flow angle at the rotor outlet. Results illustrate that a size increase (from 1.2% span to 2.3% span) in uniform clearance layout will bring adverse effects. The specific degradation in isentropic efficiency is 1.11% and 1.77% at the design and near stall points, respectively. The nonaxisymmetric rotor tip clearance arrangement significantly increases the stall margin by 25%–31% relative to the uniform layout, and the benefit is more pronounced in the case of a larger average clearance size of 48%–57%. However, there are no visible changes in the global performance characteristics. The zigzag-distributed clearance layout has a better performance impact than the other two linearly arranged schemes. Detailed analyses based on the streamwise and spanwise efficiency deviation were carried out to interpret the aerodynamic improvement. Moreover, the differentiated distribution of aerodynamics and specific flow behaviors at the rotor tip were investigated in detail over each rotor passage. Circumferentially averaged performance and radial profiles were performed. Because of the nonaxisymmetric layout, the inlet mass flow and incidence were redistributed in both radial and circumferential directions, further causing a significant alleviation of the blockage arising from the tip leakage flow, especially the region at the first 0.5 axial chord and above 80% span. The loss caused by the nonaxisymmetric tip clearance included not only the leakage flow loss, but also the interference between the leakage flow and the incoming main flow, or the low-energy fluid in the corner region of the casing, and even the loss on the suction surface of the adjacent rotor blade.

In order to further study the possibility and potentiality of hybrid electric propulsion systems applied to helicopters, this paper aims at designing and analyzing power-split hybrid electric propulsion systems (PS-HEPSs) for the AC311 helicopter equipped with a diesel engine, and investigates the fuel consumption potential of PS-HEPSs. For helicopter operating conditions, the operating characteristics of power-split transmission systems are analyzed. It is found that the operating state (power split and power cycle) of PS-HEPSs depends only on the engine speed. Based on that, the paper proposes a three-parameter matching method of transmission ratios with an objective of maximum power-split speed range. Subsequently, by establishing a mathematical model of the PS-HEPSs for helicopters and adopting dynamic programming (DP) as the energy management, the fuel consumption and working points of power sources are calculated under the specific mission and also compared with the fuel economy of parallel hybrid electric propulsion systems (P-HEPSs). Results show that, by using the proposed method, the input power-split hybrid electric propulsion system and output power-split hybrid electric propulsion system (OPS-HEPS) have similar fuel consumption, but both consume less fuel than P-HEPSs. In terms of residual power, the OPS-HEPS still has 150 Kw left, because the generator can perform as a drive motor to provide torque. Although the power-split hybrid electric systems were first applied for automobiles, the comprehensive analysis shows that the PS-HEPSs are more suitable for helicopters because the constant speed of the output shaft reduces control complexity for PS-HEPSs.

The numerical solution model of the labyrinth seal considering the multiple frequencies elliptical whirling of the rotor and swirl brake was established by the computational fluid dynamics (CFD) technique. Then, the dynamic characteristics and response force characteristics of the labyrinth seal under the condition of a preswirl inlet were studied on the basis of experimental verification of the accuracy of the numerical solution method. The influence of the circumferential solidity of the swirl brake and the preswirl ratio on the flow field between the swirl brakes and the average circumferential velocity of seal cavities were analyzed. These results show that with the increasing rotor speed, the airflow circumferential velocity in the labyrinth seal chamber is increased, and the logarithmic decrement rate of the seal-rotor system is decreased, resulting in the rotor stability being decreased. With the increasing inlet/outlet pressure ratio, the airflow circumferential velocity of the labyrinth seal cavities is decreased, the logarithmic decrement rate of the seal-rotor system is increased, causing the rotor stability is increased. With the increasing circumferential solidity of the swirl brake, cross-coupled stiffness decreases, the direct damping and effective damping increase. It contributes to rotor whirling be inhibited, and rotor stability is increased. The rotor stability can be further improved with the increasing radial length of the swirl brake. The effective damping of the labyrinth seal with 40-extended swirl brakes increases by 33.7% compared to the labyrinth seal with no swirl brakes. With the increasing the preswirl ratio, the cross-coupled stiffness increases, the direct damping and effective damping decreases, and the phase angle of the response force is decreased. It arouses that the stability of the rotor is decreased. With the increasing of whirling frequency, the cross-coupling stiffness and direct damping increase, while the effective damping decreases first and then increases, resulting in the rotor stability being decreased first and then increased. The previous research provides a theoretical basis for the structure design of the swirl brake.

Abstract

Swept blades have been widely used in the transonic fan/compressor of aircraft engines. However, because of the gap in knowledge regarding the change of the inlet flowfield by sweeping, design deviations are often produced. This study used experimental and numerical simulation methods to investigate the inlet circumferential fluctuations (CF) generated by the sweep. The results indicate that the CF source terms induced by the inviscid blade force cause a negative pressure gradient ahead of the leading edge of the forward sweep blade. This pressure gradient results in a spanwise migration of the inlet flow from the hub to the shroud, which redistributes the axial velocity across the entire spanwise range. As a result, there is a reduction in the incidence angle at the tip of the forward-swept blade, whereas an increase is observed at the root. This effect increases with increasing sweep angle. The influence of the CF stress terms starts at approximately 30% of the chord length upstream of the leading edge and the values of the axial component $UU$ and tangential component $WW$ are maximum. CF source terms are physical quantities that quantitatively describe the effect of the sweep, which can help quantify the mechanism of sweep and can play an important role in the design of sweep blades.

Practical Applications

Swept blades are extensively used in modern transonic fans and compressors to reconcile the conflict between high efficiency, high load, high through-flow, and wide stable operating range. However, there is more qualitative understanding of the mechanism of swept blades and less quantitative design theory. In this work, a method for quantitatively describing the flow mechanism of swept blades was proposed, using a reduced-dimension spanwise force balance analysis of experimental and numerical results. The effect of swept blades on the incoming flowfield is simplified as the action of the circumferential fluctuation source terms. The source terms induce a spanwise migration of the incoming flowfield, which results in a spanwise redistribution of the incidence angle of the swept blades, even under uniform inflow conditions. The circumferential fluctuation source term force can aid in quantitatively understanding the mechanism of swept blades, and enhance the precision and efficiency of swept blade design.

Through the assessments of stability derivatives to indirectly determine the structural health of transport aircraft, a method based on static aeroelasticity theory is presented. The flight data of two twin-jet transport aircraft encountering severe clear-air turbulence at transonic flight will be the study cases. One of these two twin-jet transport aircraft is an aging one and the other is a new one. The nonlinear unsteady aerodynamic models are established through flight data mining and fuzzy logic modeling techniques based on flight data. The pitch damping in oscillatory motion, controllability in longitudinal aerodynamics, and static aeroelastic characteristics in terms of the first and second aerodynamic derivatives are then estimated by using these aerodynamic models. A comparative analysis of static aeroelastic behaviors in severe clear-air turbulence for these two twin-jet transport aircraft will be assessed. The static aeroelastic effect of aging transport aircraft is more significant than that of the new one and has a larger impact on the empennage structures. The root cause of the larger static aeroelastic effects and irregular deviations of stabilizer with obvious problems of screw jack assembly will be discussed. This method can be used to assist certificate holders of aircraft maintenance factories to monitor the variations of static aeroelastic behaviors and issues of flight safety warnings for aging transport aircraft. This is used as a complementary tool for the structural maintenance program to improve aviation safety, operation, and operational efficiency.

Abstract

In aerospace load simulation loading systems, due to the coupling between the loading system and the system under testing, surplus forces and poor loading accuracies remain to be key issues. To solve this, this paper introduces a new intelligent dynamic force loading algorithm. First, based on the mechanical geometry and fluid theory of the system, a nonlinear mathematical model that can describe the system more accurately is established. By combining the mechanical structure and mathematical model of the system, it is found that the surplus force caused by external disturbances has the greatest impact on the force loading accuracy. The causes of the surplus force are then analyzed and the solutions provided. To improve the force loading performance and eliminate excess force on the controller, an intelligent dynamic force loading algorithm is proposed. This new control algorithm ensures that the surplus force caused by an external disturbance can be suppressed to within the required range in a short time as well as improve the controller performance and dynamic force loading accuracy of the system. Finally, the algorithm is applied to an aerospace air rudder load simulator and its performance is compared with the three other classical control algorithms. The results show that the proposed intelligent control algorithm can attain better accuracy in dynamic force loadings and can attain a control accuracy that meets technical requirements.

Practical Applications

The load simulator and digital controller involved in this paper are applied in a Chinese Aerospace Research Institute. The load simulator is currently used for testing and product acceptance of the institute’s servo products. Because the control algorithm proposed in this paper meets the requirements of technical indicators and product performance testing requirements, it has been recognized by the institute and has been put into practical use. The technical experts of the institute have verified the experimental results and believe that its control accuracy and loading performance are better than those of the original old load simulator.