Journal of Aerospace Engineering

Nonuniform residual stress inevitably exists in thick asymmetric component because of fabrication history originating from raw material processing. Bulk residual stress distribution of component has an impact on large-scale assembly deformation and consequently affects dimensional and shape accuracy of assembly. This paper illustrates a simulated method of introducing a component’s bulk stress state with corresponding overall deformation after machining as input to assembly analysis and decoupling the assembly deformation induced by residual stress factor through comparison quantitatively. The differences of theoretical formulating and physical meaning are discussed with the traditional method of influence coefficient (MIC) and assembly thermal expansion issue, respectively. The main procedure of potential bulk residual stress quantification for a component’s material is described involving surface stress measurement and numerical finite element analysis (FEA) processing based on normalized stress distribution shape and regression relation of key stress points. The comprehensive methodology is tested in validation experiments. The results show that the prediction compared with observation is acceptable within a certain range of error in view of uncertainties. A practical engineering case, the rear spar structure composed of edge components in horizontal stabilizer assembly, highlights the necessity of incorporating residual stress factor during further assembly analysis. The investigations of this research enhance the understanding of the effect of residual stress and help systematically improve compliant assembly precision in aircraft assembly.

Fiber reinforced polymer (FRP) composite bars are suitable substitutes for steel reinforcement in structural concrete because of their light weight and high specific strength and stiffness. Sustained load on FRP composites over the entire service life of the structure can cause failure by creep rupture at loads much smaller than those corresponding to static transient loads. Creep rupture test methods used by several researchers are presented, and a suitable loading arrangement and test method were selected for this study. The creep rupture properties of basalt FRP bars were determined using sustained loading applied to the test specimens simultaneously with an alkaline solution at an elevated temperature. The tests were based on the method specified by the American Concrete Institute (ACI) in a standard. The ultimate creep rupture strength coefficients for basalt FRP bars were determined to be approximately 18% for 50 year service life and 28% for 5 year service life. The corresponding one million hour creep coefficient was estimated to be 13%. One standard specified that creep rupture stress limits are constant regardless of the duration of the service life of the structure. Such an approach leads to conservative designs, particularly for structures with short service lives. It is recommended that time-dependent creep rupture stress limits be used in the design of structures with short service life so as to utilize the strength of FRP bars more effectively. Design creep rupture stress limits for basalt FRP bars are suggested based on the test results of this study and varying safety factors that are dependent on the duration of service life.

Prediction of thermal behavior and trajectory plays an important role in design and operation of stratospheric balloons. First, a more rounded and improved thermal and vertical trajectory model for performance prediction of stratospheric super-pressure balloons is established, including thermal models for solar radiation, infrared radiation, heat convection, and vertical dynamic model. Meanwhile, mathematical models for gas expulsion and ballast drop are also established. The new model is suitable for both the ascent stage and floating condition of stratospheric super-pressure balloon. Then, a computer program is developed based on the improved model, and predicting accuracy of the model is verified using experimental results of a stratospheric balloon. In the end, performance parameters of a new concept stratospheric balloon in ascent stage and floating condition are obtained, including film temperature, helium gas temperature, pressure differential, altitude, and velocity. Simulation results show that stratospheric balloons experience supercooling and superheating in the ascent stage and floating condition respectively, and both helium gas temperature and difference pressure have great change from day to night in the floating condition.

This study is concerned with the thrust regulation and safety protection problem for aeroengines, and a state-based switching control strategy is proposed. Specifically, the multiloop switching control framework of aeroengines is first presented, and the framework is based on switching logic and bumpless transfer. Then for the key technologies, a switching logic design that can determine which candidate controller should be put in feedback loop with the plant to achieve a desired regulation or protection performance is given. Moreover, an extended bumpless transfer design is developed for the switched system with time-varying delay and external disturbance, which can ensure that the aeroengine multiloop system performs a smooth transition at the subcontroller transfer moment. Finally, the proposed switching control strategy is applied to an aeroengine control model. As a case study, the simulations are performed to show an excellent thrust regulation and over-temperature protection performance.

A numerical investigation was undertaken to evaluate the capabilities of a differential spoiler setting (DSS) in modifying the spanwise wing load and reducing the hazard of wake vortices imposed on a following aircraft. Numerical simulations were carried out on a scaled, half-model aircraft under a high lift configuration (HLC) condition for two outboard loading cases and two inboard loading cases in addition to the baseline configuration case. The study has covered the near-wake field and the extended near-wake field regions. The results are presented in terms of velocity contours, vorticity contours, tangential velocity, circulation, and streamlines at several measurement sections downstream, namely, at $x/b=0.345$, $x/b=0.444$, $x/b=0.658$, $x/b=1.794$, and $x/b=4.484$ measured from the wing-tip trailing edge. Comparison between numerical simulations and experimental results has shown good agreement in terms of lift coefficients for all cases and fairly acceptable agreement in terms of maximum cross-flow velocity, maximum normalized total circulation, and maximum-induced rolling moment. However, in some cases the agreement is not good. Evaluation of the results has shown promising outcomes for the inboard loading cases compared with the outboard loading cases, which agree with the experimental results. The interaction between spoilers’ wake and the outboard flap tip vortex of the inboard loading cases reduces the magnitude of the tangential velocity and circulation of the wing tip vortex, delays vortex merging to a farther distance downstream, and thus, minimizes the induced rolling moment imposed on the following aircraft. Therefore, DSS can be considered a promising technique in wake vortex alleviation that can be used commercially in the future. The only limitation of this study is the need for more sophisticated computational resources to benefit from using finer mesh, which can minimize the difference between numerical and experimental results. This should therefore enable extending the study of wake vortices to the very far-wake field or up to a distance where the following aircraft is usually located.

In the design of the power system of a high-altitude airship (HAA), the principal target is to satisfy the needs of the power demand and achieve the best balance of minimum weight and maximum reliability. To handle this problem, configuration analysis of the power system is performed. Mathematical models of output power, reliability, and weight are presented. Relationships between weight, reliability, and configurations are discussed in detail. Several design rules related to the design of the HAA’s power system are deduced. For obtaining the optimal configuration, the self-adaptive genetic algorithm (GA) is applied. Results show that the optimal configuration, compared to the configuration without redundancy, has a 258% increase in reliability and a 55.9% increase in weight. The weight increase is necessary to achieve more reliability improvement.

A fault tolerant control scheme is proposed for the attitude stabilization of a small satellite using single gimbal control moment gyros (SGCMGs) and magnetic torquers (MTQs). The dynamic model of a rigid spacecraft using SGCMGs and MTQs is established, and the faults of SGCMGs are modeled as additive and multiplicative unknown dynamics. By designing a well-tuned linear extended state observer (LESO), the dynamic uncertainties, external disturbances, and the unknown faults of SGCMGs are estimated accurately and compensated actively in real time without knowing the exact spacecraft model. Then, based on the singular value decomposition (SVD) theory, the command torque is decomposed and distributed to SGCMGs and MTQs, respectively; therefore SGCMGs only need to output the torque component orthogonal to the singular direction when encountering singularity, ensuring the controllability of SGCMGs. Through comparison with a conventional proportional-derivative (PD) control method, the numerical simulations demonstrate that this scheme can tolerate potential SGCMGs faults and resolve the singularity problem, maintaining the desirable stability and performance of the satellite.

The extraction of 18 flutter derivatives of bridge decks from three degree-of-freedom (3-DOF) free vibration data using a novel direct approach is addressed in this study. Different with many conventional methods that construct a system state matrix, this approach directly extracts 18 flutter derivatives using the aeroelastically modified modal parameters. No state matrix is concerned, and thus it is more straightforward from the physical essence viewpoint. The validity and accuracy are demonstrated by a 3-DOF numerical example for bridge deck model. Afterward, the 18 flutter derivatives of two exact bridge decks with representative streamlined and bluff sections are extracted. Detailed deterministic and stochastic vulnerability analyses on identification accuracy of modal parameters and flutter derivatives were conducted for the numerical model and two bridge decks. For the free vibration method, the potential uncertainties in aeroelastic parameter determination are investigated, and the causes of low accuracy of some flutter derivatives (e.g., $H_4^{*}$, $A_4^{*}$, $P_{1\sim 4}^{*}$) are attributed to the imperfection of the linear mathematical model, testing technique, and constraint conditions, and inherent low participation and/or coupling intensities of aeroelastic components. The aeroelastic characteristics and the influence of complex aerodynamic coupling on flutter performance of both streamlined and bluff bridge decks are examined and compared to unveil the mechanisms of two kinds of flutter.

The space-based chromotomographic experiment (CTEx), a hyperspectral imager, is currently in development at the Air Force Institute of Technology. This paper details an investigation of hermetic enclosures to house commercial off-the-shelf (COTS) components. These enclosures will enable the use of electronics in space which may not be available in a space-qualified form for years and reduce cost/schedule constraints. This activity produced an experimentally validated thermal mathematical model supporting further trade-space refinement and operational planning aspects for this device. Results support the transition of this next-generation technology from the laboratory to a fully-realized, space-readied platform uniquely capable of generating hyperspectral data at high spectral and temporal resolutions.

Characteristics of wind pressures on a low-rise building immersed in various building area densities were investigated using three-dimensional Reynolds-Averaged Navier-Stokes (RANS) turbulence models. Wind pressure coefficients obtained from the numerical simulations were compared with those of wind pressure measurements. The results show that for wind pressure coefficients for an isolated model, more reasonable results were obtained when the renormalization group (RNG) $k\text{-}\epsilon$ model was used, but for a target model in a large group, there was little difference between wind pressure coefficients obtained from different turbulence models and the effect of different incident flows on wind pressure coefficients was found to be small. Lastly, the effect of area density and building distance was separately examined, showing that the effect of an upstream building was larger than that of any other buildings and area densities.

To research the performance of a solid fuel scramjet combustor, a solid fuel regression rate model was coupled into a quasi-one-dimensional flow equation, and so the flow parameters and fuel regression rate could be calculated together. According to the regression rate and the combustor diameter in the previous moment, the combustor diameter of the next moment could be obtained. The unsteady combustion and flow matter in the combustor was simplified into a steady calculation of every moment by solving the boundary condition at different moments independently. The numerical results were compared to actual experimental data from the literature, and they agreed well with the experiment. Taking the fuel/air ratio and Mach number as the optimization conditions, the initial size of the combustor could be obtained. On the basis of an optimized combustor, the relevant parameters’ variation rules were calculated and analyzed. It was found that during the working process, the combustor flow field could be kept at the supersonic flow state, the axial fuel regression rate at a specific moment first increases and then tends to be steady, the local fuel regression rate at a specific position decreases gradually with time, the mass flow rate at the combustor exit and the fuel/air ratio is relatively steady, and the total pressure recovery coefficient and the specific impulse of the combustor decline gradually.

This paper addresses dynamic modeling and performs optimization of a passive multistrut vibration isolation platform (VIP) in vibration isolation for an assembly of control moment gyroscopes (CMGs). Each strut of the VIP is composed of a metal rod and a three-parameter isolator in series. First, the frequency response function (FRF)–based substructuring method was developed to establish the dynamic model of a whole satellite system. This method can take into account the influence of the flexibility of the satellite bus and the CMG assembly as well as the gyroscopic effects of CMGs. Then, an optimization model was developed, and the FRF-based sensitivity analysis was employed to optimize the characteristic parameters of a strut under two objective functions. Finally, a numerical model was developed to verify the validity and effectiveness of the proposed method. The dynamic characteristics of an octostrut VIP were analyzed. The isolation performance was greatly enhanced after optimization. This study shows that the method is an effective and efficient tool that can be used in the design phase of VIPs.

The permeability of thermal protection blankets is measured by flowing air and helium at pressure differentials of up to 33.3 kPa across the material. Flow in both the viscous and slip regime is demonstrated by varying the average absolute pressures in the experiments from 0.400 to 80.0 kPa. Analysis shows the flow through the fibrous blanket is well characterized by a compressible formulation of Darcy’s law with the addition of the Klinkenberg slip factor. The permeability and slip parameter are obtained using multiple linear regression. A comparison between air and helium shows that the slip parameter scales with the square root of the molecular weight. Correlations between the permeability and slip factor for fibrous materials are proposed. The correlations are similar in form with previous correlations developed for rock samples, but the coefficients differ. The measured data show that the capillary radius calculated from permeability is the appropriate length scale for the Knudsen number.

In this paper, a sliding mode reconfigurable control algorithm without dedicated fault detection and identification module is developed for nonlinear aircraft systems with partial loss fault or total failure such as stuck or floating. The redundant actuators are integrated with the regular actuators in the controller seamlessly when faults or failures occur in the regular actuators. This method monitors the sliding surface to decide if the redundant actuator should be activated. The stability of the control system was proved by using Lyapunov method, and the effectiveness of the control system is validated by the simulation results of longitudinal control of a nonlinear model of Boeing 747.

In this paper, an integral feedback and feed-forward (IFF) optimal controller with hard terminal constraints for missile formation reconfiguration is designed. The controller has quadric optimal performance for expected terminal errors, output, and control quantity. From the viewpoint of the kinematics relationship of a formation in the relative coordinate frame, the authors establish a precisely linearized relative motion model by transforming the control variables. This relative motion model can intuitively manifest the relationship of the relative motion in three directions in the relative coordinate frame. In order to solve the designed IFF optimal controller, detailed deductions for deriving the related Lagrange parameters are presented. A precise integration algorithm was adopted instead of using a traditional backward integration algorithm to calculate more precise solutions for the relevant parameters in the IFF optimal controller. A collision avoidance system with four spherical domains was proposed, and a modifying principle to avoid collision during formation reconfiguration was presented. Simulation results demonstrate that the presented IFF optimal controller is capable of implementing missile formation reconfiguration rapidly, stably, and accurately. It can additionally restrain invariant or slowly varying perturbations induced by the velocity of a leader missile. Furthermore, the collision avoidance system developed in this paper can enable missiles to avoid collision during formation reconfiguration.

Determining the process noise covariance of the unscented Kalman filter (UKF) is a difficult procedure. The analytical approximation method gives satisfactory results in certain cases, but it fails when generalized for the estimation of the extended states, such as the case that sensor biases or scale factors are included in the state vector. The main aim of this research is to find an appropriate tuning algorithm for the process noise covariance of the UKF when the magnetometer biases are estimated, as well as attitude and gyro biases. In this sense, an adaptive tuning method for an UKF that is used for satellite attitude estimation is given and the adaptive UKF algorithm is tested in various scenarios for the attitude and sensor bias estimation. The given adaptation method is an easy way of tuning the filter, especially in the absence of any analytical approximation for the calculation of the process noise covariance, and the performed simulations show that by using the adaptive UKF, it is possible to get accurate estimates that are close to optimal.