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An analytical solution for spacecraft rendezvous within closed Keplerian orbits is presented using continuous radial thrust. The two spacecraft are assumed to be initially placed within the same closed orbit with arbitrary phase angle separation. The target spacecraft is assumed to always remain in the same initial orbit. The chaser vehicle, however, uses a judiciously designed maneuver sequence comprising constant-acceleration radial thrust phases interspersed with coast phases. Importantly, the proposed maneuver design is analytically characterized by one or two (depending on whether the parking orbit is circular or elliptical, respectively) nonlinear algebraic equations that determine the time intervals for the different phases. The set of initial conditions from which rendezvous is feasible is also characterized. Additionally, orbit rotation (that is, changing the orbit argument of periapsis in plane) is also demonstrated for the elliptical parking orbit case. Finally, the remarkable emergence of the golden and silver ratios, as well as their connections to Kepler’s triangle, is demonstrated for a specific sequence of thrust and coast arcs.

For optimal decision making and collaborative task performance in multispacecraft missions, it is essential for each spacecraft to maintain a state estimate of all the spacecraft in the formation. However, especially in the case of large formation sizes, each spacecraft may not be able to track or communicate with all other spacecraft. Further, for formations deployed in deep space, the unavailability of the Global Navigation Satellite System makes inertial state estimation challenging. We propose the Distributed Absolute and Relative Estimation (DARE) algorithm for autonomous inertial estimation of spacecraft formations. The algorithm enables each spacecraft to maintain an accurate inertial estimate of the entire formation even in the presence of observability and communication constraints. Each spacecraft utilizes measurements of feature points (landmarks) for inertial localization and measurements of neighboring spacecraft for relative localization. The algorithm is distributed and scalable to any number of spacecraft in the formation. A modified version of the algorithm called the Sparse Distributed Absolute and Relative Estimation (SDARE) algorithm is also derived. This algorithm is computationally more efficient at the expense of estimation accuracy, making it suitable for implementation on nanosatellites where resources are limited. Numerical simulation results demonstrating the effectiveness and comparing the performance of both algorithms are provided.

A sparse recurrent fuzzy neural network (SRFNN)-based dynamic flight envelope protection (FEP) method is proposed and integrated into an aircraft control system against possible failures. As a baseline of the aircraft control system, a fault-tolerant flight controller is used to compensate the uncertainties caused by failures, such as damaged wing. As an augmentation, the SRFNN-based dynamic FEP is employed to prevent damaged aircraft from breaking through the safe flight envelope bounds, which are updated adaptively with the help of the safety-critical information obtained by the SRFNN. Additionally, the performance of the fault-tolerant flight controller can be improved after the proposed dynamic FEP is integrated into the aircraft control system. The integrated aircraft control system not only can improve the performance of aircraft with some failures but also can prevent the damaged aircraft from getting out of control. Simulations are conducted to test and verify the effectiveness of the fault-tolerant flight controller and the SRFNN-based dynamic FEP. Simulation results show the efficacy of the proposed integrated aircraft control system.

Even in multidisciplinary aircraft design, most aspects of nonlinear controllability are only considered in later design phases. Hence, control considerations are not used to inform the overall design of the airframe. This paper presents a control-aware controller-agnostic design optimization framework that integrates nonlinear controllability constraints, which are informed by the solutions to optimal control problems, into the multidisciplinary design process. To implement the optimization, procedures to evaluate the sensitivity of the controllability constraint to variations in the design variables are defined. A supersonic aircraft case study illustrates the methodology and the steps involved in the process. The proposed approach enables the integration of control-related constraints into gradient-based design optimization.

Improved attitude control laws based on a feedback reshaping strategy using rotation matrices are proposed for attitude control of a single rigid body on [Formula: see text] and for multibody attitude synchronization on [Formula: see text]. For attitude control of a single body the proposed control law improves the closed-loop response for initial principal rotation angles near 180° while producing a globally continuous control torque. For the problem of multibody attitude synchronization, an alternative feedback reshaping strategy is proposed, which destabilizes nonconsensus equilibria and produces attitude synchronization with almost global stability and globally continuous control. Finally, a previously proposed bipartite consensus algorithm based on the concept of a structurally balanced graph with antagonistic interactions represented by negative edge weights is applied for the first time to rigid-body attitude consensus control with feedback reshaping in which the consensus attitudes and angular velocities are divided into two distinct groups. Simulations serve as illustrative examples of each of these problems and confirm the advantages of the proposed strategies. For multibody synchronization the focus is on ring graph topologies for which locally stable nonconsensus equilibria arise when using an existing attitude consensus control protocol in the literature.

This work investigates the identification of reduced-order state-space models from aeroelastic simulations of the interaction of a built-in structural finite element model and a linear aerodynamic model using unsteady potential theory. The objective is to propose and compare different new frequency-based identification methods operating on frequency response associated to the inputs and outputs of interest. The first methods considered are the Loewner interpolation method and a subspace algorithm. Subsequently, the paper introduces the possibility to apply stability constraints and to impose a certain number of poles estimated beforehand by the [Formula: see text] method in order to make the identified models closer to the true aeroelastic physics. To achieve this goal, new dedicated techniques are developed and subsequently validated on data generated by random state-space models of various orders, and on aeroelastic data obtained from structural and aerodynamic aircraft models.

This paper investigates the suitability of using a gimbaled rotor actuator for trajectory control of a projectile. In this novel concept, a single-gimbal variable-speed control moment gyro is used as the sole control actuation system, allowing actuation in two orthogonal directions and two-dimensional control over the projectile’s flight path without any moving aerodynamic surfaces. A complete nonlinear model of the eight-degree-of-freedom rotational and translational flight mechanics is developed, as well as an accurate aerodynamic model for short-range trajectories. Linear models are derived for control design, and controllers are synthesized along with a steering logic to achieve control of the projectile’s roll and sideslip angles. Successful tracking of both angles is demonstrated, and a large course deviation is achieved, showing significant maneuverability and proving the viability of the concept.

This research develops and analyzes an approach to solve the continuous monitoring problem. The analysis focuses on the string stability of a system of airspeed heterogeneous nonlinear aircraft that are controlled to overfly an endpoint at specified time intervals. The system is modeled as a cascaded set of subsystems where the position of each subsystem, an individual aircraft, is based off the estimated overfly time of the preceding vehicle. The model also includes the effects of steady-state wind and turbulence. The system is made to be string stable by nondimensionalizing a portion of the dynamics, selecting variable command limits, and solving for exponentially stable controller gains. The simulation results confirm the stability guarantees of the theoretical analysis and show that a randomized group of 20 airspeed heterogeneous aircraft experiencing realistic wind and turbulence results in an estimated arrival time error that converges toward a relatively small steady-state range.

The finite Fourier series shape-based approach for fast initial trajectory design satisfies the equations of motion and other problem constraints at discrete points by using a nonlinear programming solver to design the Fourier coefficients. In this paper, the finite Fourier series approach is extended to account for the necessary conditions for optimality during the Fourier coefficient design. The proposed method uses the approximate trajectories to analytically estimate the costates at these discrete points by employing a finite difference technique on the adjoint differential equations. Lagrange multipliers associated with any terminal state constraints are determined using an auxiliary function. Residual errors in the stationarity condition are reduced through the objective function, with the Legendre–Clebsch condition enforced as a nonlinear inequality constraint; applying these two conditions finds suboptimal solutions without an excessive loss of computational efficiency. Test cases use steering angle and thrust acceleration magnitude controls and span Earth-escape spirals, geostationary spacecraft rendezvous, and phasing maneuvers. The presented results demonstrate the proposed method’s ability to achieve trajectories closer to the optimal solution.

Liquid fuel is the choice for volume-limited propulsion systems, including detonation-based propulsion. A liquid fuel with high vapor pressure has the advantage of more fuel vapor in the mixture, which supports the transition from deflagration to detonation. This paper reports on an experimental study of deflagration-to-detonation transition (DDT) in a pulse detonation engine with heterogenous mixtures of oxygen and ethanol or acetone. Single-cycle tests were taken for different fuels, equivalence ratios, and DDT enhancement methods. The size distribution of fuel droplets was characterized at the atomizer and engine exit. The effect of the fuel evaporation was dominant for the acetone spray only. Comparing the measured detonation velocities of the two mixtures, a lower velocity deficit relative to the theoretical Chapman–Jouguet detonation velocity was measured for the acetone–oxygen mixtures, and this behavior is related to the higher amount of fuel vapor that existed in the mixtures. Moreover, a shorter transition to detonation was observed in the acetone–oxygen mixture. The addition of a Shchelkin spiral reduced the DDT distance; however, the Chapman–Jouguet condition could be reached only downstream of the obstacle. The measured detonation cell size of the heterogeneous acetone–oxygen mixture was smaller than that of the ethanol–oxygen mixture, indicating that it is more detonable.

This paper presents a computationally efficient optimization algorithm for solving nonconvex optimal control problems that involve discrete logic constraints. Traditional solution methods require binary variables and mixed-integer programming (MIP), which is prohibitively slow and computationally expensive. This paper proposes a faster and computationally cheaper algorithm that can produce locally optimal solutions in seconds. This is achieved by blending sequential convex programming and numerical continuation into a single iterative solution process. The algorithm approximates discrete logic constraints with smooth functions and uses a homotopy parameter to control the accuracy of this approximation. The homotopy parameter is updated such that, by the time the algorithm converges, the smooth approximations enforce the exact discrete logic. The effectiveness of this approach is numerically demonstrated for a realistic rendezvous scenario inspired by the Apollo Transposition and Docking maneuver. In less than 15 s of cumulative solver time, the algorithm finds a fuel-minimizing trajectory that obeys the following discrete logic constraints: thruster minimum impulse-bit, range-triggered approach cone, and range- triggered plume impingement. The optimized trajectory uses significantly less fuel than reported NASA design targets.

Despite all technological advances, runway excursions (REs) remain among the most frequent aviation accidents. Several factors have been shown to contribute to RE incidents. Although different studies have addressed the prediction of REs using quantitative data, the present study contributes to RE research by applying advanced quantitative analysis methods using text mining of voluntarily submitted reports to predict aircraft damage as the outcome of a runway excursion event. The study also intended to unveil relevant predictors from the voluntary safety reports data. The random forest model presented better results than the naïve Bayes and the gradient boosting approaches. The positive predictive value of the random forest method was measured at 74.90%, which is in contrast to 63.58% of the naïve Bayes method and 72.56% of the gradient boosting method. Although several generalizable topics were identified in the reports (such as runway and touchdown zone characteristics, flight control during approach, aircraft malfunction, weather factors, and runway surface and braking action conditions), no defined predictors for runway excursion events resulting in aircraft damage could be established.

The neural encoding by biological sensors of flying insects, which prefilters stimulus data before sending it to the central nervous system in the form of voltage spikes, enables sensing capabilities that are computationally low-cost while also being highly robust to noise. This process, which can be modeled as the composition of a linear moving average filter and a nonlinear decision function, inspired the work reported here to improve engineered sensing performance by maximizing the observability of particular neural-inspired composite measurement functions. We first present a tool to determine the observability of a linear system with measurement delay (the first element of the composition), then use a Lie algebraic observability approach to study nonlinear autonomous systems with output delay (the second element of the composition). The Lie algebraic tools are then extended to address overall observability of systems with composite outputs as in the neural encoder model we adopt. The analytical outcomes are supported using the empirical observability Gramian, and optimal sensor placement on a bioinspired wing model is performed using metrics based on the empirical Gramian., 18 pages, 7 figures

The ESA/JAXA (Japan Aerospace Exploration Agency) BepiColombo mission is currently in its cruise phase and is set to reach Mercury in late 2025. The spacecraft is equipped with advanced radio-tracking instrumentation that provides accurate radiometric data for precise orbit determination (POD). During the cruise phase, the radio link enables tests of general relativity (GR) and measurements of the solar corona properties. To fully exploit the accuracy of the radiometric measurements (Doppler and ranging) and to obtain better accuracies in the GR tests, a comprehensive dynamical model of the spacecraft is needed. A reliable and precise model of the non-gravitational accelerations would maximize the scientific return of the radio-science experiments and prevent unmodeled dynamic perturbations from degrading the POD solution. In this work, we investigate the effects of non-conservative forces acting on BepiColombo during three different radio-science campaigns conducted in November 2020, March 2021, and February 2022. The last two periods correspond to the first two GR tests of BepiColombo. We design a method for modeling these forces using telemetry measurements, radiometric observables, and mathematical models; and we analyze their characteristics in relation to the different environments encountered by the spacecraft during the three periods. This work is a preparatory and unavoidable step for the data analysis of the first two GR experiments of BepiColombo and the next radio-science campaigns, which will be performed in an even more challenging dynamical environment.

This paper presents a stability model that can make a rapid prediction of the rotating stall inception in turbomachinery and provide the spatial distribution of the corresponding instability mode. In addition, this model can take the three-dimensional geometry of blades and complex flow details in the compressor into consideration, and the solution of the development process of small perturbations can be converted to a nonlinear eigenvalue problem. We propose a solution method by converting the nonlinear eigenvalue problem into a generalized one; then, it can be solved by the Arnoldi algorithm. The proposed method can shorten the elapsed time from hundreds of hours to a few minutes, as compared with the methods adopted in previous works, substantially reducing the computational cost. Furthermore, the spatial distribution of eigenvectors can be obtained to investigate the characteristics of the perturbation mode, which can be applied as a foundation to set the inlet/outlet boundary conditions and select the eigenvalue representing the rotating stall inception. In the cases of a transonic isolated rotor and a subsonic one-stage compressor, the results are in accordance with those measured in experiments, verifying the accuracy and effectiveness of the stability model. Therefore, the model can be applied to evaluate the flow stability in the design stage of compressors with low computational cost.

This study presents an aeroelastic wind-tunnel experiment to identify the influence of the wing stiffness and hinge release threshold on the gust load alleviation performance of a folding wingtip design. Five models with different stiffness and tailoring properties are tested, and the wing root bending moment at different conditions is compared to the response with the locked-hinge condition to assess the impact on the gust load alleviation capabilities of the folding wingtip. The results show that the structural properties do not have an important impact on the peak load alleviation but the hinge release threshold and timing do. Releasing with the correct timing can reduce significantly the peak loads. However, the dynamics of the system are affected by this release; the flutter speed is decreased, and, although the performance can improve, load oscillations increase, which can be considered detrimental for reasons such as fatigue or passenger comfort.

Boundary-layer instability on a rotating cone induces coherent spiral vortices that are linked to the onset of laminar–turbulent transition. This type of transition is relevant to several aerospace systems with rotating components, e.g., aeroengine nose cones. Because a variety of options exist for the nose-cone shapes, it is important to know how their shape affects the boundary-layer transition phenomena. This study investigates the effect of varying cone angle on the boundary-layer instability on rotating cones facing axial inflow. It is found that increasing cone angle has a stabilizing effect on the boundary layer over rotating cones in axial inflow. The parameter space of Reynolds number [Formula: see text] and local rotational speed ratio [Formula: see text] is experimentally explored to find the spiral vortex growth on rotating cones of half angle [Formula: see text], 45°, and 50°. The previously addressed cases of [Formula: see text] and 30° are also revisited. Increasing half-cone angle is found to have a stabilizing effect on the boundary layer on the rotating cones with [Formula: see text]; i.e., the spiral vortex growth is delayed to higher [Formula: see text] and [Formula: see text]. This effect diminishes when the half-cone angle increases from [Formula: see text] to 50°. The spiral vortex angle [Formula: see text] decreases with increasing rotational speed ratio [Formula: see text] for all the investigated cones, irrespective of the half-cone angle. However, the instability on the broader cones is found to induce shorter azimuthal wavelengths.

Aspects of transition mechanisms on a 14° sharp-nosed ogive cylinder at Mach 6 are elucidated by considering linear and nonlinear disturbance evolution of freestream stochastic and wave packet forcing at different locations downstream of the ogive-cylinder junction. The spectral response of the stochastic forcing displays favorable agreement with experimental observations from Air Force Research Laboratory on which the configuration is based. Intermittent wave packets are generated from small-amplitude continuous freestream pressure forcing. Linear wave packet evolution reveals that Mack modes are the dominant primary instabilities, followed by relatively weaker first-mode waves. The nonlinear wavepacket displays a three-legged wall pressure perturbation signature, which is traced to spanwise curvature effects, and fundamental resonance is the dominant secondary instability mechanism. At higher amplitudes of excitation, weak signatures of subharmonic and oblique resonance phenomena are identified. Downstream placement of the small-amplitude disturbance diminishes amplification of first-mode waves; however, with high-amplitude forcing, the location of actuation does not significantly vary the disturbance evolution signature or the secondary instability mechanisms.

This paper proposes a broadband acoustic modal identification method based on nonsynchronous measurements from a microphone array, which overcomes the acoustic modal aliasing problem caused by a limited number of microphones. In the proposed method, the number of measurement points is augmented by sequentially rotating a circular microphone array that is circumferentially mounted in the wall of the compressor duct. This technique employs only one reference microphone to enhance the robustness of the cross-spectral matrix completion for nonsynchronous measurements. The fast iterative shrinkage threshold algorithm is adopted to iteratively complete the cross-spectral matrix of all measurement points for acoustic modal identification. The in-duct acoustic modal coefficients of the broadband noise are then determined using the least-squares method with the sound pressure vector recovered from the completed cross-spectral matrix. The accuracy of the cross-spectral matrix and the identified acoustic modals is verified through numerical simulations. The proposed method is validated by experimental results from an axial flow compressor. The results show that it can effectively suppress the modal aliasing problem and identify the azimuthal modes of compressor broadband noise.

Lightweight high-aspect-ratio wings make aircraft more energy efficient thanks to their lower induced drag. Because such wings exhibit large deflections, design optimization based on linear flutter analysis of the wing undeformed shape is inadequate. To address this issue, we develop a framework for integrating a geometrically nonlinear flutter constraint, which considers in-flight deflections, into a high-fidelity gradient-based structural optimization. The wing mass and stress constraints are evaluated on a linear built-up (detailed) finite element model to capture realistic structural features. The geometrically nonlinear flutter constraint is based on a condensed low-order beam representation of the built-up finite element model, which captures the impact of in-flight deflections with tractable computational effort for optimization. The flutter constraint is differentiated with respect to the detailed structural model sizing variables using the adjoint method to enable large-scale optimizations. The framework is demonstrated by minimizing the mass of a wingbox model subject to the geometrically nonlinear flutter constraint along with stress and adjacency constraints. The geometrically nonlinear flutter constraint adds a penalty of up to 60% of the baseline mass due to the impact of in-flight deflections on the flutter onset speed and its mechanism. In contrast, a linear flutter constraint evaluated on the undeformed wing adds a mass penalty of only 10%. This methodology can help design energy-efficient aircraft with high-aspect-ratio wings, which require geometrically nonlinear flutter analyses early in the design cycle.

This paper presents an overview of the system identification techniques developed and applied at the Flight Test and Research Institute during the last decade. System identification techniques have been used extensively to obtain aerodynamic models for a flight-test simulator, which was developed in house and specifically tailored for flight-test instruction and preparation for actual test flights. The methods applied were mostly the output error for data compatibility check and equation error in the time domain for model structure selection and parameter estimation. The most relevant applications are summarized in this paper, including identification of a high angle of attack model with compressibility effects for a jet trainer; estimation of drag polars at high subsonic regime for a jet fighter; identification of aerodynamic model in real-time for a propeller-driven fighter; application of flight-path reconstruction techniques to spin tests, in order to correct errors in the air data sensors; and identification of power effects on the longitudinal stability of a propeller-driven trainer. The examples show that system identification has become crucial to 1) obtain accurate aerodynamic models for training in the flight-test simulator and 2) develop more efficient flight-test techniques.

As a high-fidelity approach to transition prediction, the coupled Reynolds-averaged Navier–Stokes (RANS) and linear stability theory (LST)-based [Formula: see text] method is widely used in engineering applications and is the preferred method for laminar flow optimization. However, the further development of gradient-based laminar flow wing optimization schemes is hindered by a lack of efficient and accurate derivative computation methods for LST-based eigenvalue problems with a large number of design variables. To address this deficiency and to compute the derivatives in the LST-based solution solver, we apply the adjoint method and analytical reverse algorithm differentiation (RAD), which scale well with the number of inputs. The core of this paper is the computation of the standard eigenvalue and eigenvector derivatives for the LST problem, which involves a complex matrix. We develop an adjoint method to compute these derivatives, and we couple this method with RAD to reduce computational costs. In addition, we incorporate the LST-based partial derivatives into the laminar–turbulent transition prediction framework for the computation of total derivatives. We verify our proposed method with reference to finite difference (FD) results for an infinite swept wing. Both the intermediate derivatives from the transition module and total derivatives agree with the FD reference results to at least three digits, demonstrating the accuracy of our proposed approach. The fully adjoint and the coupled adjoint–RAD methods both have considerable advantages in terms of computational efficiency compared with iterative RAD and FD methods. The LST-based transition method and the proposed method for efficient and accurate derivative computations have prospects for wide application to laminar flow optimization in aerodynamic design.

In the present work, we discuss the employment of a computational fluid dynamics approach to evaluate the specific impulse of solid rocket motors. Particular care is focused on two-phase flow and divergence losses, which represent the most important contributions to the overall nozzle performance loss. A comprehensive parametric study is performed on the Zefiro 9A nozzle with the aim to evaluate the detrimental influence of relevant key features, such as alumina particle dimension, polydispersion, crystallization, and motor operating conditions. The capability of the present model to represent, with good accuracy, the overall performance of solid rocket motors is demonstrated by comparing the experimental specific impulse of several motors with numerical predictions.

This paper summarizes several key issues of importance related to intercept of high-speed maneuvering targets under stratospheric altitude conditions, where it is unlikely that the interceptor will have a significant speed and/or maneuverability advantage. It explains why one cannot rely on traditional proportional navigation guidance and presents simulation results for a recently developed finite time method of guidance. A simplified version of finite time intercept guidance is derived, and a modification required to allow it to be practically implemented is described. Simulation results are presented to illustrate the advantages of the simplified version.

Pneumatic probes such as five-hole probes (5HP) can conveniently measure three-dimensional flow angles, plus total and static pressure. In most applications, transducers are connected using pneumatic tubes, allowing the probe head to be highly miniaturized and robust. However, such “steady” probes are often used in unsteady flows, where they measure a pneumatically averaged flowfield that can differ from the time mean. To better understand these pneumatic averaging effects, an analytical framework is constructed using a quasi-steady model. Total and static pressure coefficients have a symmetric response to both positive and negative incidence. When incidence fluctuates, there is therefore a bias in the pneumatic average. These errors are evident in a shedding wake experiment, where a 5HP overestimates total pressure loss by up to 44% compared to a Kiel probe. These effects can be predicted by coupling an unsteady Reynolds-averaged Navier–Stokes calculation with the quasi-steady model. By predicting pneumatic averaging errors, the quasi-steady model can be used to obtain like-for-like validation of calculations against experimental data. Measurement data can also be corrected, provided that flow angle fluctuations can be measured or estimated. This approach can be readily used to postcorrect the large body of historical data likely to have been corrupted by pneumatic-averaging errors.

We present a novel stochastic optimal control framework that accounts for various types of uncertainties, with application to reentry trajectory planning. The formulation of the optimal trajectory control problem is presented in the context of an indirect method where a functional objective associated with the terminal vehicle speed is to be minimized. Uncertain input parameters in the optimal trajectory control model, including aerodynamic parameters and initial and terminal conditions, are modeled as aleatory random variables, while the statistical parameters of these aleatory distributions are themselves random variables. The parametric and model uncertainties are simultaneously propagated through an extended polynomial chaos expansion (EPCE) formalism. Several metrics are described to evaluate response statistics and presented as insightful tools for robust decision making. Specifically, the response probability density function (PDF) reflecting influence of both epistemic and aleatory uncertainties is obtained. By sampling over the random variables representing model error, an ensemble of response PDFs is generated and the associated failure probability is estimated as a random variable with its own polynomial chaos expansion. Besides, the sensitivity index functions of response PDF with respect to the statistical parameters are evaluated. Coupling parametric and model uncertainties within the EPCE framework leads to a robust and efficient paradigm for multilevel uncertainty propagation and PDF characterization in general optimal control problems.

At transitional Reynolds numbers, an elastically supported airfoil oscillating in pitch can undergo laminar separation flutter (LSF), which is characterized by self-sustained small-amplitude oscillations. To gain insight into the mechanism of LSF, we conduct wind tunnel tests for [Formula: see text] to investigate the LSF response of a freely rotating NACA0015 airfoil with various structural natural frequencies and positions of the elastic axis. The experimental results show that the dominant mode of flow around the NACA0015 airfoil abruptly changes at [Formula: see text], resulting in a change in the trend of LSF response. Then, an aeroelastic model is constructed to explain how the instability of LSF arises. This model can accurately predict the LSF frequency of airfoils with various structural natural frequencies. Moreover, based on the aeroelastic model, we perform linear stability analysis on the aeroelastic system and propose the instability criterion for LSF as [Formula: see text]. This instability criterion is identical to that for rotational galloping of square cylinder, indicating that LSF and rotational galloping of square cylinder are essentially the same aeroelastic phenomenon that appears in different aerodynamic profiles. This finding identifies the underlying cause for the remarkable similarities between the two phenomena.

Debris strikes on operational spacecraft are becoming more common due to increasing numbers of space objects. Sample return missions indicate hundreds of minor strikes, but rigorous analysis is often only performed when a strike causes an anomaly in spacecraft performance. Developing techniques to identify and assess minor strikes that do not immediately cause anomalous behavior can help to validate models for debris populations and aid in the attribution of future anomalies. This study develops methods to detect subtle abrupt orbit perturbations indicative of minor debris strikes. An extended Kalman filter with dynamic model compensation is used to estimate a spacecraft’s orbit state based on simulated full-state (i.e., GPS) measurements. The filter is applied to the data forward and backward in time, and then a modified Fraser–Potter smoother is used to produce a fused state estimate. Various test statistics are developed and compared to identify abrupt unexpected changes in spacecraft velocity; techniques include McReynold’s filter-smoother consistency test and the Mahalanobis distance between forward and backward filter states. A trade study is performed to investigate the performance of test statistics as a function of filter parameters, and a Monte Carlo analysis illustrates the filter’s ability to detect and estimate strikes.

Dual quaternions can describe both three-dimensional translation and three-dimensional rotation and have the advantages of a simple form, accurate modeling, and precise physical meaning in terms of dynamic modeling. The traditional rigid-body dynamics modeling theory based on dual quaternions is only applicable to the case of modeling for the center of mass. When modeling appendages, the flexible appendages must be symmetrically configured to reduce the modeling complexity. However, in space operations, flexible appendages are usually not symmetrically installed, which significantly limits the applications of dual-quaternion theory. In this paper, we propose a new dual-quaternion derivative rule and a rigid-body dynamic modeling method suitable for any point. The theory can be applied to model flexible appendages with any configuration without changing the mathematical modeling framework. Finally, we demonstrate the integrated dynamic equation of a rigid–flexible coupling system considering the coupling of position and attitude, and we verify the correctness of the model by numerical simulation.

This paper discusses a flush air-data sensing (FADS) system for a sharp-nosed hypersonic vehicle designed to estimate flight air data in real time at hypersonic speeds. The design’s target condition is Mach 5.0 to 7.0 with an angle of attack within 5 deg. The FADS system estimates air data by relating them to surface pressures measured from surface-mounted ports on the vehicle forebody. Unique combinations of pressure ratios were used to estimate the freestream dynamic pressure, which was a primary parameter for the flight experiment, the angle of attack, and the Mach number. The proposed FADS estimation algorithm was validated through numerical simulation, which was also used to generate datasets of surface pressures for given flight conditions. To handle possible sensor errors related to estimation accuracy in real-time estimation for a flight experiment, redundant systems were implemented. The results indicated that the designed FADS system can estimate air data within the uncertainty of 4.8% for a single estimator by considering sensor errors for freestream dynamic pressures in the range of 0–100 kPa, including targeted and off-design flight conditions. The proposed algorithm can estimate the air data with an acceptable level of uncertainty while retaining the robustness of estimation to sensor failures with low computational cost.

The present experimental investigation characterizes the development and decay of transonic buffet on the supercritical airfoil OAT15A, as well as the flow field and shock motion throughout the fully developed buffet cycle. A comprehensive database of high-speed focusing schlieren sequences provides access to the temporal and spectral behavior of the oscillatory shock motion within a broad parameter space, covering Mach numbers 0.68–0.80 and angles of attack 3.3–6.5 deg. From these data, multiple temporal and spectral properties characterizing the distinctness of the shock oscillation are extracted. Combining these indicators enables us to clearly distinguish fully established buffet from pre-buffet, buffet onset, and buffet offset. This criterion helps to interpret the diverse findings in the relevant literature cases. Particle-image velocimetry measurements provide detailed insight into the turbulence structure and behavior throughout a fully established buffet cycle at a Mach number of 0.72 and an angle of attack of 5 deg. Instantaneous and phase-averaged velocity maps indicate massive flow separation for shock positions close to the upstream limit. Velocity profiles illustrate the large intermittent velocity deficit persisting until the trailing edge and corroborate most pronounced separation when induced by an upstream-traveling shock wave. Representative phases of the buffet cycle visualize and quantify the severe modification of the flow topology modulated by the shock and lay the groundwork to improve the understanding of mechanisms governing shock buffet.

Grid assembly is one of the key components of an ion thruster and directly affects the performance and life of the thruster. The measurement of the sputtering yield of the grid assembly under ion beam bombardment is highly significant for predicting the lifetime of the grid assembly and ion thruster. This study systematically summarizes the main methods currently used for sputtering yield measurement of grid materials, analyzes the advantages and disadvantages of different measurement methods, and provides suggestions for sputtering yield measurements in the low-energy ([Formula: see text]) range. In addition, this study compares the sputtering resistance properties of metal- and carbon-based grid materials and summarizes the influence of key core parameters, such as the surface roughness, surface morphology, binding energy, and incident angle, on the sputtering yield. The results can be used to guide the correction of the sputter yield theoretical formula and the numerical simulation of sputtering erosion.

A coaxial low-energy surface flashover (LESF) ignitor for CubeSat electric propulsion systems was developed and tested. The ignitor features a coaxial geometry with copper electrodes directly bonded to the inner and outer surfaces of the alumina ceramic tubular insulator. The ignitor proved to be operational throughout (and after) an extended duration test of 10 million pulses. Characterization of a single LESF event via intensified charge-coupled device fast photography showed that the initial plasma was generated along the insulator surface, while the later plasma production was governed by the column attached to the copper electrodes. The plasma plume propagated primarily perpendicular to the insulator surface at around [Formula: see text]. Further investigation on the erosion of ceramic insulator and copper electrodes via energy-dispersive x-ray spectroscopy analysis of a witness plate exposed to LESF and scanning electron microscopy observation of the electrodes revealed that the ceramic erosion ([Formula: see text] molecules per pulse) was predominant over electrodes erosion ([Formula: see text] atoms per pulse or [Formula: see text]).

This study estimates the probability of extreme fatalities from civil airplane accidents in the United States since the early 1980s. To this end, extreme civil airplane accidents were modeled using extreme value statistics, allowing for changes in the distribution due to other factors. The analysis revealed that an extreme aircraft accident causes 18 to 20 fatalities every two years with costs between 111.6 and 124 million U.S. dollars (USD); whereas accidents with (at least) between 70 and 100 fatalities will occur every 20 years, costing between 424 and 620 million USD. The findings also indicate that higher volumes of air traffic increase the return levels of severe aviation accidents, whereas higher ticket prices reduce the likelihood of these. The findings emphasize the serious consequences of airplane accidents with respect to the number of fatalities but also how these depend on market conditions.

Origami-based structures have expanded in recent years due to new mathematical formulations along with materials that can achieve the bending requirements, often in the form of composites. While current methods of manufacturing can produce complex structures, they lack the ability to scale efficiently. A novel manufacturing technique is discussed in this work that allows for a simpler and lower cost fabrication that can scale to larger structures through robotic deposition. Samples made from this technique are investigated to understand the mechanical bending performance and effect on the tensile properties. Results show an orientation-dependent response for the material with the 45° samples having a direct impact on the tensile response. However, their bending response proved to be stiffer compared to the [Formula: see text] samples, holding more consistent bend radii. Joint stacking was also investigated, where discrete layers were not bonded together and showed an increase in force required to bend compared to the completely bonded samples. The results provide insight into how integrated composite hinges can perform in complex structures. The advancement of composite origami technology additionally works to reduce the overall number or parts and fasteners that are needed to achieve detailed deployable structures.

Two-temperature models for coupled vibrational relaxation and dissociation in shock-heated oxygen are assessed using low-uncertainty measured data from reflected shock tube experiments. A computationally efficient multistep technique is developed to model the unsteady dynamics of shock reflection in a relaxing and dissociating gas. The developed technique is then benchmarked through comparison with unsteady computational fluid dynamic simulations. Results from the benchmarking effort demonstrate that the adopted multistep modeling procedure accurately captures the dominant gas dynamic effects influencing the state of the test gas at the measurement location. A parametric study is then performed to assess several combinations of possible two-temperature modeling approaches for nonequilibrium oxygen dissociation. The current assessment demonstrates that the widely adopted Park model is inconsistent with the measured data, while the recently developed modified Marrone and Treanor (MMT) model demonstrates promising agreement with the data. The results of the present study clearly indicate that the MMT model is more appropriate for two-temperature modeling of nonequilibrium oxygen dissociation than the legacy Park model. Patterns in the parametric comparison also suggest that the approximate treatment of non-Boltzmann vibrational state distributions within the MMT model may require improvement.