Journal of Spacecraft and Rockets

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ISSN / EISSN : 0022-4650 / 1533-6794
Total articles ≅ 9,904
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Karin W. Fulford, Dale Ferguson, Ryan Hoffmann, Vanessa Murray, Daniel Engelhart, Elena Plis
Journal of Spacecraft and Rockets pp 1-5;

GPS satellites undergo surface contamination on the solar array coverglasses from repeated arcing events. Using NASA Air Force Spacecraft Charging Analyzer Program (Nascap-2 K) spacecraft charging simulation software, a GPS Block IIF satellite model was constructed and analyzed in realistic Medium Earth Orbit environments. GPS Block IIF satellites have Qioptiq CMG-type coverglasses (as do all other GPS satellites). The Nascap-2 K model with CMG coverglasses charges to high differential levels in maximum charging environments, in the range above the arcing threshold, as determined by studies at the Air Force Research Laboratory (AFRL), and so arcing is confirmed by theory. This finding agrees with onboard Los Alamos National Laboratory measurements and Arecibo observational data for GPS satellites. Other AFRL results show that CMX-type coverglasses, being more bulk-conductive, should charge less and perhaps mitigate arcing on the solar arrays. A Nascap-2 K model using CMX coverglasses is shown to charge differentially much less than CMG, and not reach levels above the arcing threshold. In the simulation, the commonly used CMG coverglass charges quickly, exceeding its arcing voltage threshold of 1500 V in about 1 h and 10 min. In comparison, CMX results indicate an ability to remain well under its arcing threshold throughout the orbit.
Ken Fujii, Akiko Matsuo, Junichi Oki, Hideyuki Taguchi, Takahiro Chiga, Yutaka Ikeda
Journal of Spacecraft and Rockets pp 1-11;

This paper examines thermal response behind the high-Mach integrated control (HIMICO) experiment’s engine. In this experimental aircraft, instruments must be protected from heat load due to exhaust gas; therefore, a coupling calculation between the fluid and the wall is conducted to confirm the performance of HIMICO’s thermal protection system (TPS). First, the validity of the coupling calculation is confirmed through comparison with aerodynamic heating on a hollow cylinder. The present calculation result can reproduce the surface temperature distribution on the cylinder better than previous work has managed because we consider the turbulence effect. Second, one-dimensional heat-transfer analysis is conducted on the external nozzle, and the appropriateness of the calculation result is confirmed through comparison with the ramjet-engine experiment. Finally, a coupling calculation between the fluid and the wall is conducted to investigate the local temperature distribution. The calculation result indicates that the temperature increase easily meets design requirements and that TPS performance is sufficient.
Chenyu Lu, Zhitan Zhou, Xiaoyang Liang, Guigao Le
Journal of Spacecraft and Rockets pp 1-8;

This paper investigates the thermal environment of launch pads during rocket takeoff. The rocket plume model is set up using the three-dimensional Navier–Stokes equations and realizable k−ε turbulence model. The aluminium oxide particles in the plume are considered by the Eulerian dispersed phase model. Comparing with the experimental data, the accuracy of the numerical model can be verified. On this basis, in total 16 numerical cases are performed to analyze the influence of the rocket flight altitude and lateral drift on the temperature of the launch pad. The results show that the launch pad has a more severe thermal environment at the flight altitudes from 3 to 20 m. Because of the rocket drift, a high-temperature gas flow layer is formed on the launch pad, which results in a dramatic increase in the temperature of the cross beam. At this location, the maximum temperature can reach up to 3900 K, 30% higher than the value without rocket drift. The method in this paper can provide an effective way to evaluate the thermal environment of the launch pad during rocket launching and have a great value on the thermal protection system design.
Tadayoshi Shoyama, Yutaka Wada, Takafumi Matsui
Journal of Spacecraft and Rockets pp 1-9;

A hybrid rockoon, which is a hybrid rocket launched from a high-altitude balloon, is proposed. The rocket system configuration was studied for single-stage or multiple-stage rockets, and the performance and range safety requirements were considered. The propellants of the hybrid-rocket motor are a low-melting-point thermoplastic fuel and nitrous oxide, which are beneficial, owing to their mechanical properties at cold temperatures experienced in high-altitude environments. Three-dimensional launch trajectory analyses were performed for science missions aimed at sampling cosmic dust levitating in the upper stratosphere. The apogee altitude can be significantly increased by elevating the altitude of the launch point because the problem of low thrust level of the hybrid rocket is solved by increasing the nozzle expansion ratio. Suborbital trajectories with multiple apogee points and orbital missions to extremely low Earth orbits are presented, which provide a long flight path in the upper atmosphere. The sequence of events and flight characteristics of the proposed hybrid rockoon system are discussed, and necessary technological improvements in the structure and propulsion systems are presented.
Eugene P. Bonfiglio, Mark Wallace, Eric Gustafson, Min-Kun Chung, Evgeniy Sklyanskiy, Devin Kipp
Journal of Spacecraft and Rockets pp 1-10;

Early in operational testing for the InSight mission to Mars, it was discovered that the final maneuver to target the entry-interface point was unexpectedly sensitive to planned atmosphere updates that would be based on real-time measurements of the Martian atmosphere by Mars Reconnaissance Orbiter. Upon investigation, the team realized that the Phoenix mission also discovered this sensitivity during operational testing. Further investigation identified that the sensitivity was a result of the fact that both the entry flight path angle and ground target were being held fixed in the maneuver design process, constraining the maneuver in a way that forced the entry time to change to compensate for changes to the nominal trajectory from updating the atmosphere model. The final maneuver was planned for 22 h before entry, at which point it is very expensive to change entry time. The study also revealed that any unguided Mars entry, descent, and landing mission would be impacted by this sensitivity if it used real-time atmosphere observations to model the nominal expected atmosphere used for maneuver targeting of the entry-interface point. This paper discusses the results of that investigation and presents a number of mitigations as well as the consequences of ignoring the sensitivity.
Ananthalakshmy K. Moorthy, John J. Blandino, Michael A. Demetriou, Nikolaos A. Gatsonis
Journal of Spacecraft and Rockets pp 1-17;

A wide variety of missions could be enabled by extended orbital flight in extreme low Earth orbit, defined as an altitude range of 150–250 km. This study investigates the feasibility of a nanosatellite (mass <10 kg) using a propulsive, attitude control system in conjunction with a primary propulsion system to extend mission life. The primary propulsion system consists of a pair of electrospray thrusters providing a combined thrust of 0.12 mN at 1 W. Pulsed plasma thrusters are used for attitude control. The mission consists of two phases. In Phase I, a 4U CubeSat is deployed from a 414 km orbit and uses the primary propulsion system to deorbit to an initial altitude within the targeted range of 244±10 km. Phase I lasts 12.73 days, with the propulsion system consuming 5.6 g of propellant to deliver a ΔV of 28.12 m/s. In Phase II the mission is maintained until the remaining 25.2 g of propellant is consumed. Phase II lasts for 30.27 days, corresponding to a ΔV of 57.22 m/s with a mean altitude of 244 km. Using this approach, a primary mission life of 30.27 days could be achieved, compared with 3.1 days without primary propulsion.
Mohammad Abbas, David W. Riggins, Michael D. Watson
Journal of Spacecraft and Rockets pp 1-14;

The physical basis and application of the fundamental relationship governing the balance and utilization of available energy for a chemical rocket operating within the atmosphere are described. The relative contributions of the thermochemical availability and the kinetic energy of the stored propellant to the overall energy availability are shown. There are optimal flight velocities for which 1) overall entropy generation is minimized and 2) effectiveness of the conversion of available energy to vehicle force power is maximized. The fundamental impacts of entropy generation within a rocket engine flowfield on energy utilization and thrust/performance characteristics of a rocket-powered vehicle are studied analytically. Highly nonlinear coupling between entropy generation in the engine and entropy generation in the wake is observed; this is true as well as for other energy utilization parameters, such as thrust losses. Representative energy utilization-based performance maps for selected (example) legacy rocket systems across altitude/flight velocity confirm the theoretical results. Propulsion models and available flight data are then used to provide the time evolution of energy utilization characteristics for the flight of Apollo 11 (Saturn V).
, Osamu Mori, Masanori Matsushita, Nobukatsu Okuizumi, Yasutaka Satou, Junichiro Kawaguchi
Journal of Spacecraft and Rockets pp 1-17;

A novel approach for shape control of membrane structures is presented to realize their use in three-dimensional and variable configurations. The shape control is accomplished by exciting a spinning membrane. The membrane forms a shape consisting of several vibration modes, depending on the input frequency, and the wave surface stands still when its frequency is synchronized with the spin rate; that is, the wave propagation and the spin cancel each other, resulting in a static wave surface in the inertial frame. This idea enables control of continuous membrane structures with large deformation using fewer actuators than conventional methods. This paper describes the general theory of the static wave-based shape control. The mathematical model of membrane vibration, the classification of control input, and the control system for exciting a static wave are summarized. The proposed method is demonstrated through a ground experiment. A 1 m large polyimide film is rotated and vibrated in a vacuum chamber, and the output shape is measured using a real-time depth sensor. It is shown that the observed shapes agree with numerical simulation results. An additional simulation that models the Japanese solar sail Interplanetary Kite-craft Accelerated by Radiation Of the Sun (IKAROS) demonstrates that the proposed method also works with a practically large-scale membrane in the space environment.
Li Si Ye, Sun Zhen Sheng, Quan Hui, Zhang Cui Ping
Journal of Spacecraft and Rockets pp 1-17;

Aiming at the problem of silo design, the silo ejector model and ejector function were proposed. The properties of the static pressure matching function were studied. The critical conditions were analyzed when the engine total pressure, the diameter of the silo, and the outlet pressure of the mixing chamber changed. It is proved that the curve that shows the inlet pressure of the mixing chamber changes along with the outlet pressure of the mixing chamber is a straight line parallel to the horizontal axis when the outlet pressure of the mixing chamber is smaller than the stagnation critical point, the curve that shows the inlet pressure of the mixing chamber changes along with the total pressure of the nozzle develops along the optimal pressure function curve when the total pressure of the nozzle is greater than the stagnation critical point, and the curve that shows the inlet pressure of the mixing chamber changes along with the diameter of the mixing chamber develops along the optimal pressure function curve when the diameter of the mixing chamber is smaller than the stagnation critical point. The characteristic curves of the silo ejector and the ideal ejector were compared when imbalance or wall friction were of concern, respectively, which show that the degree of imbalance has little effect on the calculation results of the silo ejector and the wall friction has a significant effect, especially near the stagnation point. The results provide important guidance for the design of silos and ejectors. Finally, the reliability of the ejector function method is verified by comparing the experimental data with the theoretical results.
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