Journal of Spacecraft and Rockets

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ISSN / EISSN : 0022-4650 / 1533-6794
Total articles ≅ 9,814
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Hiroumi Tani, Tetsufumi Ohmaru
Journal of Spacecraft and Rockets pp 1-14; https://doi.org/10.2514/1.a34882

Abstract:
The present study performs hybrid Navier–Stokes (N-S) and direct simulation Monte Carlo (DSMC) simulations to predict surface pressures and heat fluxes imposed by plumes impinging on spacecraft in actual length scale. The first application was the impingement of H-II Transfer Vehicle main engine plumes on the International Space Station. It was found that the profiles of the surface pressure and heat flux imposed by the plume interactions strongly depend on the high-density regions generated by the interactions among multiple plumes and between plumes and upstream structures. The plume of a lunar lander was another application. A plume impinging on the lunar surface generated a strong shock wave near ground level, and some of the hot gases imposed larger pressures and heat fluxes as the altitude decreased. The third application was a cold-gas plume vented from rocket tanks. Although the nozzles were small, plumes expanded rapidly and subjected some rocket assemblies to various forces and cooling effects. Last, the last application to a thruster plume impinging on debris showed that the plume can impart enough torque to detumble the rotation of a debris. This study adequately demonstrated that the usefulness of the hybrid N-S/DSMC technique for assessing risks or exploring new ideas regarding plume impingement in actual spacecraft length scale, including the plume–plume and plume–surface interactions.
Dennis Nikitaev,
Journal of Spacecraft and Rockets pp 1-11; https://doi.org/10.2514/1.a35289

Abstract:
Recently, NASA’s Artemis Program has considered sustainably returning humans to the moon with in situ resource utilization as the central focus and considers nuclear thermal propulsion (NTP) to be a viable propulsion system candidate for deep space missions. This program also considers the moon to function as a staging area for deep space missions due to its strategic orbit around Earth. The Lunar Crater Observation and Sensing Satellite discovered the abundancy of water, ammonia, and other volatiles in permanently shadowed regions on the moon. Because NTP can theoretically use any fluid as a propellant, water and ammonia are two potential alternative propellants to hydrogen that do not require postprocessing to be used directly in situ. Alternative propellant NTP engine models were developed to examine the performance of water and ammonia. It was found that bleed cycle architectures are infeasible with current turbine materials, but expander cycles have acceptable turbine inlet temperatures with marginal performance differences from bleed cycles. At reactor power levels comparable to hydrogen NTP engines, water doubles the thrust while ammonia increases it by 70%, although at a much lower specific impulse. The largest contributor to the life of the engine was the maximum temperature of the SiC cladding for water and the maximum temperature of the fuel for ammonia where higher values resulted in shorter engine life.
Enrico Schiassi, , , Fabio Curti,
Journal of Spacecraft and Rockets pp 1-16; https://doi.org/10.2514/1.a35138

Abstract:
This paper presents a novel framework, combining the indirect method and Physics-Informed Neural Networks (PINNs), to learn optimal control actions for a series of optimal planar orbit transfer problems. According to the indirect method, the optimal control is retrieved by directly applying the Pontryagin minimum principle, which provides the first-order necessary optimality conditions. The necessary conditions result in a two-point boundary value problem (TPBVP) in the state–costate pair, constituting a system of ordinary differential equations, representing the physics constraints of the problem. More precisely, the goal is to model a neural network (NN) representation of the state–costate pair for which the residuals of the TPVBP are as close to zero as possible. This is done using PINNs, which are particular NNs where the training is driven by the problem’s physics constraints. A particular PINN method will be used, named Extreme Theory of Functional Connections (X-TFC), which is a synergy of the classic PINN and the Theory of Functional Connections. With X-TFC, the TPBVP’s boundary conditions are analytically satisfied. This avoids having unbalanced gradients during the network training. The results show the feasibility of employing PINNs to tackle this class of optimal control problems for space applications.
Mattia Pugliatti, ,
Journal of Spacecraft and Rockets pp 1-17; https://doi.org/10.2514/1.a35213

Abstract:
In this work data-driven image processing options for a CubeSat mission around a binary asteroid system are investigated. The methods considered belong to two main branches of image processing methods: centroid and artificial intelligence. The former is represented by three variations of centroiding methods, and the latter by three neural networks and one convolutional neural network. The first contribution of this work is an enhanced center of brightness method with a data-driven scattering law. This method is demonstrated to share similarities with neural networks in terms of both design and performance, with the advantage of relying on a traditional, robust, and fully explainable algorithm. The second contribution is given by the performance assessment between the different families of image processing methods. For this purpose, the Milani mission is considered as a case study: a 6U CubeSat that will visit the Didymos system as part of the Hera mission. From this analysis, it emerges that convolutional networks perform better than other methods across all metrics considered. This hints to the importance of filtering techniques to extract spatial information from images, which is a unique feature of the convolutional approach over the other image processing methods considered.
Casey R. Heidrich, Michael J. Sparapany, Michael J. Grant
Journal of Spacecraft and Rockets pp 1-15; https://doi.org/10.2514/1.a35229

Abstract:
Constraints in optimal control problems introduce challenges with traditional indirect methods. Bang-bang/singular solutions with discontinuous or indefinite control laws add further difficulty in numerical solution. Recent efforts in control regularization strategies have sought to overcome these limitations. Regularization generates a smoothed constraint transformation of a multiphase Hamiltonian boundary value problem to a single-phase unconstrained problem. This work develops a new approach to regularization using orthogonal error-control saturation functions. The method is developed for problems in bang-bang/singular form. The method is then applied to problems of general Hamiltonian structure using system extension and differential control. Applications in state constraint regularization are discussed. A key feature of the new approach is to eliminate ambiguity of the control law derived from the first-order necessary conditions of optimality. Results show desirable stability and convergence in numerical continuation. The method is applied to classical problems in optimal control, as well as problems of interest in aerospace mission design.
Jeffrey Scott Clawson,
Journal of Spacecraft and Rockets pp 1-12; https://doi.org/10.2514/1.a35178

Abstract:
A linear covariance simulation framework is developed and validated for a two-dimensional target engagement scenario. The scenario comprises a single interceptor equipped with an inertial navigation system aided by absolute position measurements, as well as range, angle, and range rate measurements relative to the incoming target. The interceptor uses a proportional navigation guidance law to engage the target, modeled as nominally constant velocity perturbed by Singer motion accelerations. The linear covariance framework is developed by linearizing the differential and measurement equations about the nominal trajectory, and forming an augmented system comprising truth and navigation state dispersions. In contrast with sample-based methods, the developed linear covariance framework can calculate the truth state dispersion covariance and the estimation error covariance throughout the engagement in a single run. It also provides several advantages over analysis methods such as the adjoint technique or traditional covariance analysis. The linear covariance computational efficiency is exploited to rapidly analyze a head-on target engagement problem.
Saima Bukhat Khan, , Suhail Akhtar, Dan Xie, Rizwan Riaz
Journal of Spacecraft and Rockets pp 1-15; https://doi.org/10.2514/1.a35211

Abstract:
A successful atmospheric entry vehicle (AEV) design demonstrates manageable aerodynamic heating, bearable structural load, smooth deceleration, and intended trajectory during a descent into the atmosphere. Interestingly, the supersonic regime of the AEV manifests limit cycle oscillations (LCO) that restrict the maneuver potential and deployment of the drag chute. In this paper, efforts are made numerically and analytically to link the causation of LCO with external geometric variables of AEV. Computational fluid dynamics is used to calculate damping derivatives. The numerical results are validated with the Orion Crew Exploration Vehicle. The multiple time scales method, which belongs to the class of perturbation methods, is explored to develop an approximate closed-form solution of the nonlinear dynamic behavior of AEV. The analytical solution identifies that higher-order nonlinearities associated with pitch damping and static lift govern the onset of LCO. Finally, a parametric interaction study is carried out to determine the effect of two design variables, apex angle and length, on the vehicle’s dynamic stability. The mean data values from the main geometric effects plot show the condition of finite-amplitude oscillations. The results indicate that variation in these two parameters significantly impacts the magnitude of identified higher-order nonlinearities.
Valentin Buyakofu, Ken Matsuoka, Koichi Matsuyama, Akira Kawasaki, Hiroaki Watanabe, Noboru Itouyama, Keisuke Goto, Kazuki Ishihara, Tomoyuki Noda, , et al.
Journal of Spacecraft and Rockets pp 1-11; https://doi.org/10.2514/1.a35200

Abstract:
This paper presents the results of an S-shaped pulse detonation engine (PDE) ground firing test in the form of a detonation engine system. The world’s first technology demonstration of PDE in space using a sounding rocket is planned, and the aim is to control the rocket spin rate in the axial direction using pulsed detonation. The PDE operation at full sequence was successful. The despin rate change of the rocket between continuous oxygen supply and successful PDE operation is expected to be 0.95  deg/s per run. This change in despin rate can be measured by an onboard gyro sensor, making the system flyable. The test results were compared with data from thrust measurement tests conducted in a laboratory, the results of which confirmed the thrust generation under an ambient pressure of 0.5±0.1  kPa . The average thrust values in the thrust measurement experiments showed good agreement of 101±3% with a quasi-steady-state model introduced to predict the PDE thrust. These results demonstrate the feasibility of the newly developed PDE and its system as the world’s first technology demonstration of detonation propulsion in space.
Robert Alviani, Devon Fano, Jonathan Poggie, Gregory Blaisdell
Journal of Spacecraft and Rockets pp 1-14; https://doi.org/10.2514/1.a35183

Abstract:
Aerothermodynamic loading in the gap between a generic-missile-shaped body and a control fin was investigated through computations employing the Reynolds-averaged Navier–Stokes equations. The computed results for fully turbulent Mach 6 flow were compared with wind tunnel data obtained at the Arnold Engineering Development Center in 1979 experiments. For each case, the computational mesh encompassed the full geometry. Baseline computations of the missile body, in the absence of a fin, showed reasonable agreement with experimental data for heat transfer, surface pressure, and pitot pressure. The computations captured the order of magnitude increase in heat transfer over baseline levels with the presence of the fin and connecting cylinder. In agreement with the experiments, the computations predicted that minimizing the gap height minimized heat transfer levels. The maximum computed heat transfer levels on the missile surface occurred very close to the cylinder, a region that was inaccessible experimentally. In addition, heat transfer levels on the windward surface of the cylinder, which were not explored in the experimental study, significantly exceeded those on the missile surface for larger gap heights. The intense aerodynamic heating observed for this configuration highlights the high demands for thermal protection in the gap region under a control fin on a high-speed vehicle.
Boris Benedikter, , Guido Colasurdo, Simone Pizzurro, Enrico Cavallini
Journal of Spacecraft and Rockets pp 1-16; https://doi.org/10.2514/1.a35194

Abstract:
This paper presents a convex programming approach to the optimization of a multistage launch vehicle ascent trajectory, from the liftoff to the payload injection into the target orbit, taking into account multiple nonconvex constraints, such as the maximum heat flux after fairing jettisoning and the splash-down of the burned-out stages. Lossless and successive convexification methods are employed to convert the problem into a sequence of convex subproblems. Virtual controls and buffer zones are included to ensure the recursive feasibility of the process, and a state-of-the-art method for updating the reference solution is implemented to filter out undesired phenomena that may hinder convergence. A hp pseudospectral discretization scheme is used to accurately capture the complex ascent and return dynamics with a limited computational effort. The convergence properties, computational efficiency, and robustness of the algorithm are discussed on the basis of numerical results. The ascent of a VEGA-like launch vehicle toward a polar orbit is used as a case study to discuss the interaction between the heat flux and splash-down constraints. Finally, a sensitivity analysis of the launch vehicle carrying capacity to different splash-down locations is presented.
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