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Robert W. Maddock, Alicia M. Dwyer Cianciolo, Daniel K. Litton, Carlie H. Zumwalt
Journal of Spacecraft and Rockets pp 1-11; https://doi.org/10.2514/1.a35022

Abstract:
On 26 November 2018, the Interior Exploration Using Seismic Investigations, Geodesy and Heat Transport (InSight) lander successfully touched down on the surface of Mars. Over its more than seven year development, NASA Langley Research Center’s Program to Optimize Simulated Trajectories II (POST2) was used to assess the mission’s entry, descent, and landing (EDL) vehicle system performance against related requirements across the full range of possible environmental and spacecraft conditions. Due to the high degree of entry body, lander platform, and EDL design heritage, much of the simulation code was derived from the 2007 Mars Phoenix Lander mission. The InSight POST2 six-degree-of-freedom simulation included models for Mars atmosphere, gravity, and digital elevation maps of the landing location. Additionally, vehicle-specific aerodynamic, parachute, engine, navigation sensor, flight software, and landing radar models were included. A set of dispersions for each model, as well as for additional simulation input parameters, was also included in order to provide a statistical, Monte Carlo prediction of the EDL system performance. An overview of the preflight performance assessments completed, including the various simulation campaigns used, will be provided. Ultimately, this work was critical in the assessment of readiness for the InSight launch. A brief description of the use of this simulation in support of flight operations is also discussed.
Casey R. Heidrich, Marcus J. Holzinger, Robert D. Braun
Journal of Spacecraft and Rockets pp 1-14; https://doi.org/10.2514/1.a35175

Abstract:
Entry flight is a critical mission phase of planetary aeroassist problems. During atmospheric flight, aleatory-epistemic uncertainty and environmental factors reduce the accuracy of predicted future states for precision targeting. This problem has been approached historically with closed-loop guidance rooted in certainty equivalence. This property separates estimation and control problems, allowing each to be considered independently. In other concept studies, an observer model is neglected altogether in favor of assuming perfect state knowledge. However, a flight system will inevitably have imprecise state information and variability in its underlying dynamics and measurement models. Systemic uncertainty is a fundamental limitation of existing entry guidance approaches. This work seeks to overcome these challenges by posing aerocapture as a robust optimization problem. The cost objective of the maneuver is reformulated to account for uncertainty in atmospheric structure, vehicle performance parameters, and state estimation accuracy using an observer-based consider filter. An expected value performance cost is developed from anticipated measurement conditioning effects. A rapid solution methodology is illustrated using explicit integration strategies with a parameterized control structure. Results for a Mars aerocapture concept study show improvement in the postcapture orbit accuracy with low computational overhead.
Gerardo Saucedo-Zárate, Marco Saucedo-González, José Luis Arauz-Lara, Ángel de la Cruz-Mendoza, Emmanuel Vázquez-Martínez, Hernán González-Aguilar, José Refugio Martínez-Mendoza, Azdrubal Lobo-Guerrero
Journal of Spacecraft and Rockets pp 1-4; https://doi.org/10.2514/1.a34999

Brian R. Hollis
Journal of Spacecraft and Rockets pp 1-24; https://doi.org/10.2514/1.a34791

Abstract:
An experimental investigation of hexcomb-pattern surface roughness effects on boundary-layer transition and convective heating has been performed. Two representative entry vehicle geometries, a spherical-cap aeroshell and a sphere-cone aeroshell, were considered. Multiple cast ceramic models of each geometry were fabricated with various roughness pattern densities and depths that simulated an ablated hexcomb-structure thermal protection system. Wind-tunnel testing was performed at Mach 6 over a range of Reynolds numbers sufficient to produce laminar, transitional, and turbulent flow. Aeroheating and boundary-layer transition onset data were obtained using global phosphor thermography. The experimental heating data are presented herein, as are comparisons to laminar and turbulent smooth-wall heat transfer distributions from computational flowfield simulations. The experimental data were used to develop a unified boundary-layer transition correlation for both sand-grain distributed roughness and hexcomb-pattern roughness.
Yih-Kanq Chen
Journal of Spacecraft and Rockets pp 1-13; https://doi.org/10.2514/1.a35167

Abstract:
Analysis of multidimensional in-depth thermal response for the phenolic impregnated carbon ablator (PICA)/CV-1144-0 material system in NASA Langley Research Center Hypersonic Materials Environmental Test System environments was performed using the 3-D Fully Implicit Ablation and Thermal (3dFIAT) code, which was enhanced to enable the simulation of multidimensional thin-coating burn-through and multistreams of pyrolysis gas flow. CV-1144-0 is a controlled volatility room-temperature vulcanized silicone oxygen protective coating. Three proposed in-depth thermal response models for CV-1144-0 were examined to assess their accuracy and understand how the coating affected PICA thermal performance. These proposed in-depth thermal response models were used to predict the in-depth temperature history for three groups of arcjet test models. Pyrometer surface temperature measurements and recessed surface data from laser scans were used as the boundary conditions in 3dFIAT in-depth response simulations. Data-parallel line relaxation was used to compute the surface heat flux and pressure distributions based on the surface geometries of pretest and posttest models. Comparisons between computations and thermocouple data are presented and discussed for each test model. The uncertainty of surface temperature and recession measurements on in-depth temperature prediction is investigated. The effect of CV-1144-0 coating thickness on the modeling of PICA in-depth thermal response is also studied.
Yusuke Oki, Hiroyuki Okamoto, Takahiro Sasaki, Toru Yamamoto, Keiichi Wada
Journal of Spacecraft and Rockets pp 1-9; https://doi.org/10.2514/1.a35055

Abstract:
This paper describes the system design of satellites with a blocking effect. The blocking effect is that in which an object interferes with the radio waves used for communication or the sunlight used for power generation. The consideration of this effect is necessary in proving the feasibility of the satellite system design, especially in various missions, such as active debris removal (ADR) and on-orbit servicing, because the satellite must approach closely and touch the noncooperative objects. This study investigated the system feasibility of an ADR satellite using the system design method by considering the blocking effect. First, we propose a feasible attitude plan that satisfies the condition of the depth of discharge of the battery, considering that the generated power is affected by the debris. Next, the power-generation distribution on the solar array paddle was investigated. Finally, the feasibility of the satellite’s communication with the ground station and GPS satellites under the blocking effect was verified.
Michael Guthrie, Michael R. Ross
Journal of Spacecraft and Rockets pp 1-15; https://doi.org/10.2514/1.a35099

Abstract:
This work explores deriving transmissibility functions for a missile from a measured location at the base of the fairing to a desired location within the payload. A pressure on the outside of the fairing and the rocket motor’s excitation creates an acceleration at a measured location and a desired location. Typically, the desired location is not measured. In fact, it is typical that the payload may change, but measured acceleration at the base of the fairing is generally similar to previous test flights. Given this knowledge, it is desired to use a finite-element model to create a transmissibility function which relates acceleration from the previous test flight’s measured location at the base of the fairing to acceleration at a location in the new payload. Four methods are explored for deriving this transmissibility, with the goal of finding an appropriate transmissibility when both the pressure and rocket motor excitation are equally present. These methods are assessed using transient results from a simple example problem, and it is found that one of the methods gives good agreement with the transient results for the full range of loads considered.
Hao Chen, Melkior Ornik, Koki Ho
Journal of Spacecraft and Rockets pp 1-14; https://doi.org/10.2514/1.a35077

Abstract:
The trend of space commercialization is changing the decision-making process for future space exploration architectures, and there is a growing need for a new decision-making framework that explicitly considers the interactions between the mission coordinator (i.e., government) and the commercial players. In response to this challenge, this paper develops a framework for space exploration and logistics decision making that considers the incentive mechanism to stimulate commercial participation in future space infrastructure development and deployment. By extending the state-of-the-art space logistics design formulations from the game-theoretic perspective, the relationship between the mission coordinator and commercial players is first analyzed, and then the formulation for the optimal architecture design and incentive mechanism in three different scenarios is derived. To demonstrate and evaluate the effectiveness of the proposed framework, a case study on lunar habitat infrastructure design and deployment is conducted. Results show how total mission demands and in-situ resource utilization system performances after deployment may impact the cooperation among stakeholders. As an outcome of this study, an incentive-based decision-making framework that can benefit both the mission coordinator and the commercial players from commercialization is derived, leading to a mutually beneficial space exploration between the government and the industry.
Andrew Jolley, Greg Cohen, Damien Joubert, Andrew Lambert
Journal of Spacecraft and Rockets pp 1-10; https://doi.org/10.2514/1.a35015

Abstract:
Broadband photometry has been used for many years to infer basic information about satellites; however, there has been limited success at remotely determining satellite surface materials, and the variability of brightness and spectral energy distribution of satellite reflections complicates satellite identification. This paper demonstrates the potential utility of event-based sensors for remotely characterizing satellites based on the reflectance properties of their surface materials. Event-based sensors offer three important advantages over traditional frame-based sensors, such as charge-coupled devices (CCDs) for characterizing satellite materials on orbit: very high temporal resolution, high dynamic range, and low data rates. This allows rapid, fine-resolution measurements over a broad range of intensities. An event-based camera was used to characterize the broadband reflectance properties of five common satellite materials over a range of illumination and observation angles in the laboratory. Some of the results are very distinctive, and have not previously been reported, demonstrating that event-based sensors might perform better than CCDs at satellite identification and material characterization. The results also show that different materials can exhibit quite different, and sometimes very distinctive, illumination/observation geometry-dependent reflectance characteristics. Using high angular resolution event-based data to assess broadband reflectance changes with changing geometry could be an effective means of unambiguously identifying satellites or determining the presence of specific satellite surface materials.
Xuze Lu, Chao Li, Sheng Zhang, Yunpeng Li, Biaosong Chen
Journal of Spacecraft and Rockets pp 1-11; https://doi.org/10.2514/1.a35101

Abstract:
Predictions of the remaining lifetime of aging solid-propellant rocket motors are affected by the uncertainties of various chemical parameters in the reaction. This paper proposes a nonprobabilistic interval method for evaluating the life bounds of aging solid-propellant rocket motors. In this method, chemical uncertainties in the aging process are represented by uncertain-but-bounded initial parameters, and the uncertainties arising in aging solid-propellant rocket motors are described by an interval model. A sparse regression Legendre inclusion function is then employed to establish an interval model for aging solid-propellant rocket motors, which can be directly solved in deterministic form using the finite element method. Finally, in conjunction with the deterministic solutions, the sparse regression Legendre interval algorithm is employed to obtain the bounds of the storage life related to the uncertain interval parameters. The effectiveness of the proposed method is verified by considering an uncertain storage life prediction problem under the accelerated thermoaging process with constant temperature conditions. The proposed method provides a reliable tool for predicting the storage life of aging solid-propellant rocket motors.
Ian G. Clark, Clara O’Farrell, Christopher D. Karlgaard
Journal of Spacecraft and Rockets pp 1-11; https://doi.org/10.2514/1.a35180

Abstract:
On 26 November 2018, the Mars InSight lander successfully touched down at Elysium Planitia and began its 26 month prime mission. The entry, descent, and landing sequence of InSight included an 11.8 m supersonically deployed disk–gap–band parachute. This paper describes the reconstructed performance of the supersonic parachute of InSight on Mars. Measurements from the onboard inertial measurement unit along with prelaunch measurements of the parachute system and spacecraft, assumptions about the aerodynamics of the vehicle, and models for the Martian atmosphere have been used to reconstruct the trajectory of the spacecraft and the performance of the parachute system. The reconstruction results were compared against preflight predictions. Reconstruction of the InSight trajectory leading up to parachute deployment showed that the vehicle trimmed in a lift-down orientation during entry, and thus experienced greater deceleration than expected by most preflight simulations. This led to parachute deployment conditions that diverged from the nominal preflight predictions. The parachute was mortar deployed at a Mach number of approximately 1.5, below the nominal preflight expectation of 1.66. The peak inflation load was 45 kN, well below the 67 kN design limit load of the parachute. Following deployment of the parachute, the rotational rates of the vehicle and the dynamics of the system were in excellent agreement with preflight expectations.
Ira Katz, Bao Hoang, Kazuhiro Toyoda
Journal of Spacecraft and Rockets pp 1-9; https://doi.org/10.2514/1.a35163

Abstract:
For decades it has been known that spacecraft charging-induced electrostatic arc discharge on solar arrays can cause cell degradation and damaging secondary arcs. While discharging on large-area solar arrays is the concern, there is no space data on primary arc current magnitudes on large arrays. Laboratory testing is normally performed on coupons with just a few solar cells. External capacitors and circuitry based on theoretical models are used to simulate the arc current profile from the charge stored on the rest of the array. Previously, the current profiles for large arrays were generated assuming that the arc plasma expanded at a constant velocity and instantaneously discharged cells at the plasma perimeter. Arc-current profiles using this perimeter model failed to agree with laboratory tests on large arrays. An improved model needs to account for the range of expansion velocities in a multicomponent plasma, and that the cell neutralization at the perimeter is not instantaneous but is limited by the arc-plasma thermal electron current density. A new, multifluid model of arc-plasma expansion addresses both concerns. The model assumes as input primary arc ion generation rates. Ion expansion velocities are calculated self-consistently from electric potentials in the expanding plasma.
, , Fanny Keller, Alessandro Zuccaro Marchi, Roger Walker
Journal of Spacecraft and Rockets pp 1-13; https://doi.org/10.2514/1.a34945

Abstract:
This paper describes the results of a system study for the in-space assembly of large optical telescopes using a building block approach involving multiple mirror segments. Such an approach is enabled by CubeSat rendezvous and docking and mirror steering technologies for compensating misalignment of the assembly. This concept thus removes the constraints imposed by the limited volume of launchers’ fairing, and hence larger telescope dimensions can be envisaged, leading to new observation capabilities. Each mirror segment is mounted on top of a small satellite, containing the miniaturized subsystems to achieve autonomous docking to either the service module containing the science instruments or other segments to build up the primary mirror. A systems’ design of all the elements constituting the assembly is provided, including tradeoffs, optimization, as well as mass and power budgets. The proposed solution scales with the size of the telescope and can be used to build primary mirrors as large as 20 m in diameter. The critical guidance, navigation, and control function is introduced, and related docking performances are presented. The results indicate that the in-space assembly of such large telescopes is feasible once these two key technologies are demonstrated in flight in the coming years.
Soumyo Dutta, Christopher D. Karlgaard, Ashley M. Korzun, Justin S. Green, Jake A. Tynis, Joseph D. Williams, Bryan Yount, Alan M. Cassell, Paul F. Wercinski
Journal of Spacecraft and Rockets pp 1-24; https://doi.org/10.2514/1.a35090

Abstract:
Sounding Rocket One (SR-1), the first flight test of the Adaptable Deployable Entry and Placement Technology (ADEPT), was performed on September 12, 2018. ADEPT is a deployable aeroshell that is stowed for launch and deployed before atmospheric flight to increase the drag area of the spacecraft. The main objectives of the SR-1 flight test were to demonstrate that the ADEPT vehicle deploys exo-atmospherically and to characterize the stability of the vehicle during atmospheric flight. The SR-1 test vehicle was a 0.7-m-diam, 70°-half-angle faceted sphere-cone and was the primary payload on an UP Aerospace Spaceloft launch vehicle from the White Sands Missile Range. ADEPT successfully separated from the spent booster in its stowed configuration, opened above 100 km altitude, and landed in the deployed configuration within White Sands Missile Range. ADEPT was able to reach peak Mach number of 3.1 and was able to show angle-of-attack stability through Mach 0.8, which was the objective of the mission. The aerodynamics and flight mechanics of the vehicle were modeled preflight for performance and range safety predictions. After the flight, the on-board instrumentation was used to reconstruct the flight performance. This paper describes the predictions and postflight reconstruction, and how the predictions compared with the flight data.
Robin A. S. Beck, Jarvis T. Songer, Christine E. Szalai, David A. Saunders, Mark A. Johnson, Chris Karlgaard
Journal of Spacecraft and Rockets pp 1-8; https://doi.org/10.2514/1.a35078

Abstract:
The Mars Interior Exploration using Seismic Investigations, Geodesy and Heat Transport (InSight) spacecraft, which successfully touched down on the planet surface on November 26, 2018, was proposed as a near build-to-print copy of the Mars Phoenix vehicle to reduce the overall cost and risk of the mission. Because the lander payload and the atmospheric entry trajectory were similar enough to those of the Phoenix mission, it was expected that the Phoenix thermal protection material thickness would be sufficient to withstand the entry heat load. However, allowances were made for increasing the heatshield thickness because the planned spacecraft arrival date coincided with the Mars dust storm season. The aftbody thermal protection system components were not expected to change. In a first for a U.S. Mars mission, the aerothermal environments for InSight included estimates of radiative heat flux to the aftbody from the wake. The combined convective and radiative heat fluxes were used to determine if the as-flown Phoenix thermal protection system design would be sufficient for InSight. Although the radiative heat fluxes on the aftbody were predicted to be comparable to, or even higher than the local convective heat fluxes, all analyses of the aftbody thermal protection system showed that the design would still be adequate. Aerothermal environments were computed for the vehicle from postflight reconstruction of the atmosphere and trajectory and compared with the design environments. These comparisons showed that the predicted as-flown conditions were less severe than the design conditions.
Sanford M. Krasner, Kristoffer Bruvold, Jared A. Call, Paul D. Fieseler, Andrew T. Klesh, M. Michael Kobayashi, Norman E. Lay, Ryan S. Lim, David D. Morabito, Kamal Oudrhiri, et al.
Journal of Spacecraft and Rockets pp 1-13; https://doi.org/10.2514/1.a34892

Abstract:
The Interior Exploration Using Seismic Investigations, Geodesy and Heat Transport (InSight) spacecraft landed successfully on 26 November 2018 to conduct an exploration of the interior of Mars. To meet NASA’s requirement for communications during critical events, the InSight lander transmitted telemetry continuously throughout entry, descent, and landing. This allowed the public to witness the landing in real time. The transmissions were received by five assets: three at Mars and two on Earth. These included real-time relay of telemetry by the first deep-space CubeSats. This paper describes the constraints on the design of an entry, design, and landing communication link; the uncertainties in the trajectory; and the modeling that was used to meet these constraints. It then reports on the actual performance of each link, including an unexpected degradation of signal to the Mars Reconnaissance Orbiter. The lessons learned from this process contributed to the successful relay of data from the Perseverance lander, as well as to the design of future landing communications systems.
Samuel W. Albert, , Robert D. Braun
Journal of Spacecraft and Rockets pp 1-14; https://doi.org/10.2514/1.a34953

Abstract:
The co-delivery of a direct-entry probe and an aerocapture orbiter from a single atmospheric entry state is a novel way to include ride-along probes or orbiters on interplanetary missions. This is made possible through combining two technologies: low-cost small satellites and aerocapture. This study investigates the feasibility of this co-delivery method from a flight-mechanics perspective. The availability of direct-entry and aerocapture trajectories from a single entry flight-path angle is assessed for a large range of feasible ballistic coefficients at Earth, Mars, Venus, Titan, and Neptune. Apoapsis altitude, peak heat flux, total heat load, and peak g-load are also quantified across this trade space. A representative scenario implementing closed-loop guidance is presented for a proof of concept, and the trajectory dispersions due to relevant uncertainties are quantified in a Monte Carlo analysis. Passive ballistic impactor or penetrator probes as a secondary mission with a primary lift-modulated aerocapture orbiter is identified as the most promising configuration.
Joanna Fulton,
Journal of Spacecraft and Rockets pp 1-20; https://doi.org/10.2514/1.a34938

Abstract:
Self-actuated deployable space structures present a novel challenge for deployment dynamics modeling efforts, where the system-level influence of strain energy components must be captured. Here, the free deployment of an origami-folded structure through actuation of strain energy hinges is studied. Studies include experimental testing, multibody dynamics modeling, and finite element modeling. An approach for modeling high strain tape spring hinges for use in a multibody simulation of free-deployment dynamics analysis is presented and demonstrated. This approach considers hinges with multiple degrees of freedom beyond the primary fold axis angle. A novel folded deployable structure is designed and prototyped with a segmented fold pattern and strain energy hinges integrated in the design. A suite of deployment tests is conducted on the prototype using videogrammetry. A full simulation of the prototype is constructed from a multibody dynamics model and the hinge model, and the predicted deployment behavior for relative hinge states is evaluated against the experimental testing. Additionally, the prototype deployment is replicated using an explicit dynamic finite element analysis for a performance comparison. The models demonstrate strong correlation for deployment time predictions across the relative hinge states, and the finite element analysis correlates all deployment behaviors.
Julian Hammerl,
Journal of Spacecraft and Rockets pp 1-11; https://doi.org/10.2514/1.a35165

Abstract:
The electrostatic tractor has been proposed to touchlessly remove space debris from geosynchronous orbit by taking advantage of intercraft Coulomb forces. A controlled spacecraft (tug) emits an electron beam onto an uncooperative or retired satellite (debris). Thus, the tug raises its own electrostatic positive potential to tens of kilovolts, whereas the debris charges negatively. This results in an attractive force called the electrostatic tractor. Prior research investigated the charged relative motion dynamics and control of the electrostatic tractor for two spherical spacecraft and how charge uncertainty affects the relative motion control stability, but attitude effects could not be studied due to the two-sphere model. This work uses the multisphere method to consider general three-dimensional spacecraft shapes, and it investigates how the electric potential uncertainty and debris attitude impact the equilibrium separation distance between the two craft. The results show bounds for safe operations that avoid a collision. State regions are identified where the relative motion is particularly sensitive to potential uncertainty. The relative station-keeping performance using either higher- or lower-fidelity multi-sphere method models are compared to demonstrate that even a lower-fidelity multi-sphere method model can yield good results.
Michael E. Holloway, Ross S. Chaudhry, Iain D. Boyd
Journal of Spacecraft and Rockets pp 1-12; https://doi.org/10.2514/1.a35052

Abstract:
The effect of thermochemical kinetics modeling on hypersonic flow over a double-cone geometry is investigated. The double-cone is simulated using three different approaches based on the Park model: nonequilibrium flow, equilibrium flow, and frozen flow for air at four different freestream conditions. The thermochemical model effects on the flowfield and surface properties are specific areas of interest. The resulting aerothermodynamic loads are compared to experiments performed in the CUBRC LENS-XX facility and indicate that thermochemistry modeling plays an important role in determining surface properties. The results indicate that the specific thermochemistry model used to describe hypersonic flow over a double-cone significantly affects surface properties for both CUBRC facilities, especially at high enthalpies. A comparison of Park and Modified Marrone–Treanor thermochemistry models is also made, and it is concluded that the models produce similar surface properties, due to the freestream density, and fail to reproduce experimental results. Consistent overprediction of the pressure drag and heating rate indicates there is some unknown fundamental difference between the experiments and the simulations, thus limiting the usefulness of these double-cone experiments for validation of thermochemistry models.
Chenyu Lu, Zhitan Zhou, Xiaoyang Liang, Guigao Le
Journal of Spacecraft and Rockets pp 1-8; https://doi.org/10.2514/1.a35128

Abstract:
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.
Ken Fujii, Akiko Matsuo, Junichi Oki, Hideyuki Taguchi, Takahiro Chiga, Yutaka Ikeda
Journal of Spacecraft and Rockets pp 1-11; https://doi.org/10.2514/1.a35018

Abstract:
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.
Karin W. Fulford, Dale Ferguson, Ryan Hoffmann, Vanessa Murray, Daniel Engelhart, Elena Plis
Journal of Spacecraft and Rockets pp 1-5; https://doi.org/10.2514/1.a35106

Abstract:
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.
Ananthalakshmy K. Moorthy, John J. Blandino, Michael A. Demetriou, Nikolaos A. Gatsonis
Journal of Spacecraft and Rockets pp 1-17; https://doi.org/10.2514/1.a34975

Abstract:
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.
Eugene P. Bonfiglio, Mark Wallace, Eric Gustafson, Min-Kun Chung, Evgeniy Sklyanskiy, Devin Kipp
Journal of Spacecraft and Rockets pp 1-10; https://doi.org/10.2514/1.a35181

Abstract:
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.
Tadayoshi Shoyama, Yutaka Wada, Takafumi Matsui
Journal of Spacecraft and Rockets pp 1-9; https://doi.org/10.2514/1.a35040

Abstract:
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.
Mohammad Abbas, David W. Riggins, Michael D. Watson
Journal of Spacecraft and Rockets pp 1-14; https://doi.org/10.2514/1.a35100

Abstract:
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).
Li Si Ye, Sun Zhen Sheng, Quan Hui, Zhang Cui Ping
Journal of Spacecraft and Rockets pp 1-17; https://doi.org/10.2514/1.a34963

Abstract:
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.
Yonatan Amit-Shapira, Pini Gurfil, Eviatar Edlerman
Journal of Spacecraft and Rockets pp 1-17; https://doi.org/10.2514/1.a35032

Abstract:
Satellite formation establishment may consume considerable fuel if cross-track distance (CTD) is a mission requirement. This work proposes a strategy for significantly reducing the required fuel. Instead of direct out-of-plane maneuvers, it is suggested to use in-plane maneuvers, while using the nodal precession caused by the J2 perturbation. The CTD dynamics are analyzed, and representative formation geometries are considered. Analytical expressions for the CTD and related establishment timing are derived for the case of impulsive thrust. For most cases, the expected CTD establishment duration is reasonable, and the fuel saving with the J2-assisted CTD establishment strategy is considerable. For the case of J2-assisted CTD establishment with continuous thrust, optimal control methods are used. The proposed strategy is proven to be optimal for two cases: limited and unlimited thrust magnitude. Finally, the cases of J2-assisted CTD establishment with impulsive and with continuous thrust are compared. It is shown that continuous thrust is preferable in terms of fuel, for similar CTD establishment durations.
, Osamu Mori, Masanori Matsushita, Nobukatsu Okuizumi, Yasutaka Satou, Junichiro Kawaguchi
Journal of Spacecraft and Rockets pp 1-17; https://doi.org/10.2514/1.a35084

Abstract:
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.
Katiyayni Balachandran, Kamesh Subbarao
Journal of Spacecraft and Rockets pp 1-19; https://doi.org/10.2514/1.a34976

Abstract:
The importance of determining the sidereal rotation period of an astronomical object on future investigations pertaining to said object has been well documented in the literature. Researchers, however, have differed in their techniques used to estimate and model objects in the space catalog. In this paper, several period-estimation methods will be explored ranging across Fourier and phase-folding techniques. These methods will be tested using ground-based observations of light curve data for various resident space objects that fall under a rigid body context (i.e., asteroids, satellites, probes, rocket bodies) and celestial objects like stars and extrasolar planets. The effect of varying sample size, the inadequacies in unevenly sampled data processing, autonomy of the method, and complexity of parameters are investigated. For the models of artificial space objects that are not open source, a simulation is used to generate synthetic light curves with which all of the above-mentioned techniques are also employed. To account for heterogeneity in method parameters, each technique is tested with a range of values to optimize the rotational period. Results for uniformly sampled asteroid data as well as nonuniformly sampled stellar objects and generic sinusoidal data show variances in accuracy of the methods, but certain methods stand out.
Pablo Machuca, Joan-Pau Sánchez
Journal of Spacecraft and Rockets pp 1-18; https://doi.org/10.2514/1.a34986

Abstract:
Recent advancements in CubeSat technology unfold new mission ideas and the opportunity to lower the cost of space exploration. Ground operations costs for interplanetary CubeSats, however, still represent a challenge toward low-cost CubeSat missions: hence, certain levels of autonomy are desirable. The feasibility of autonomous asteroid flyby missions using CubeSats is assessed here, and an effective strategy for autonomous operations is proposed. The navigation strategy is composed of observations of the Sun, visible planets, and the target asteroid, whereas the guidance strategy is composed of two optimally timed trajectory correction maneuvers. A Monte Carlo analysis is performed to understand the flyby accuracies that can be achieved by autonomous CubeSats, in consideration of errors and uncertainties in a) departure conditions, b) propulsive maneuvers, c) observations, and d) asteroid ephemerides. Flyby accuracies better than ±100 km (3σ) are found possible, and main limiting factors to autonomous missions are identified, namely a) on-board asteroid visibility time (Vlim≥11), b) ΔV for correction maneuvers (>15 m/s), c) asteroid ephemeris uncertainty (<1000 km), and d) short duration of transfer to asteroid. Ultimately, this study assesses the readiness level of current CubeSat technology to autonomously flyby near-Earth asteroids, in consideration of realistic system specifications, errors, and uncertainties.
Maximilien Berthet, Kojiro Suzuki
Journal of Spacecraft and Rockets pp 1-13; https://doi.org/10.2514/1.a34948

Abstract:
The two leading perturbations acting on satellites in the orbital decay region of low Earth orbit are gravitational nonsphericity of the Earth and atmospheric drag. A detailed understanding of their properties is required to accurately predict satellite dynamics, such as the orbital lifetime. However, there is still room to improve current perturbation models in this region, especially for the atmosphere. The purpose of this paper is 1) to develop a simple method to reverse engineer the leading perturbations from two-line elements (TLEs), and 2) to apply the method to the 2017 Re-Entry Satellite with Gossamer Aeroshell and GPS/Iridium (EGG) nanosatellite mission as a case study. Because of the low ballistic coefficient of EGG and its low orbital altitude, flight was highly affected by both aerodynamic and geopotential effects. Secular and long-periodic effects are extracted via linearized perturbation theory. The results show that EGG acted as a low-cost passive multiphysics sensor for several phenomena, including the changing solar flux, the diurnal atmospheric bulge, rotating winds, and the zonal geopotential. For example, the second zonal geopotential harmonic coefficient is estimated to within 0.008%. To the authors’ knowledge, this is the first study to produce estimates of both atmospheric and geopotential perturbations via TLEs from a single satellite.
Gilles Bailet, Amandine Denis, Alexis Bourgoing, Christophe O. Laux, Thierry E. Magin
Journal of Spacecraft and Rockets pp 1-13; https://doi.org/10.2514/1.a34968

Abstract:
Quantifying the thermal response of a heat shield is a key step in the design of a spacecraft to ensure its survivability during atmospheric entry. Recession and swelling of the thermal protection material have a drastic effect on both the heat transfer within the vehicle and aerothermodynamic transients. Postflight analysis of reentry can be achieved after recovery on Earth, but it is more difficult for entries on other bodies of the solar system. A dedicated instrumentation is necessary to understand the evolution of the thermal protection material thickness during flight. This paper investigates the current limitations of the available measurement techniques. A low-mass passive solution is proposed to measure with high accuracy the phenomena of recession and swelling. The QubeSat for Aerothermodynamic Research and Measurements on Ablation (QARMAN) CubeSat mission provides a flight opportunity to develop a dedicated payload to quantify the recession and swelling of an ablative heat shield made of P50 cork from AmorimTM. In addition, this payload allows us to study the radiation of a reentry plasma in the presence of ablation products through the same optical path. The performance of the measurement technique and the integration of the instrument are discussed for the QARMAN platform, demonstrating its applicability to a space mission.
Sandeep Soman, , Prasanth P. Nair, Heuy Dong Kim
Journal of Spacecraft and Rockets pp 1-13; https://doi.org/10.2514/1.a34992

Abstract:
A linear aerospike nozzle is a kind of altitude adaptive nozzle that has optimum performance over the entire operational range. In this work, flow characteristics of a linear plug nozzle and a truncated plug with 40% plug length are studied numerically. Different flow types associated with an increase in nozzle pressure ratio, corresponding wall pressure variation and shock patterns are studied. The effect of base pressure improvement with a secondary flow from the base (base bleed) of the truncated nozzle is also examined. Four base bleed points are considered, and the optimum bleed point is selected based on performance and base pressure improvement. The two-dimensional numerical model is solved using Reynolds-averaged Navier–Stokes equations with the Transition SST turbulence model. Validation shows that the present model is good enough at predicting the flow features associated with a linear plug nozzle. From the three different flow types associated with this nozzle, the transition from one type to other happens at lower nozzle pressure ratios with truncation due to the reduction in length of the plug. Thrust coefficient increased by 3.54% in the case of the bleed 3 configuration with a secondary flow of 2% of inlet mass flow rate when compared with 40% plug length without base bleed.
Brittanny V. Holden, Shan He,
Journal of Spacecraft and Rockets pp 1-14; https://doi.org/10.2514/1.a35014

Abstract:
The problem of minimum-time, low-thrust, Earth-to-Mars interplanetary orbital trajectory optimization is considered. The minimum-time orbital transfer problem is modeled as a four-phase optimal control problem where the four phases correspond to planetary alignment, Earth escape, heliocentric transfer, and Mars capture. The four-phase optimal control problem is then solved using a direct collocation adaptive Gaussian quadrature collocation method. The following three models are used in the study: 1) circular planetary motion, 2) elliptic planetary motion, and 3) elliptic planetary motion with gravity perturbations, where the transfer begins in a geostationary orbit and terminates in a Mars-stationary orbit. Results for all three cases are provided, and one particular case is studied in detail to show the key features of the optimal solutions. Using the particular value thrust specific force of 9.8×10−4 m⋅s−2, it was found that the minimum times for cases 1, 2, and 3 are, respectively, 215, 196, and 198 d with departure dates, respectively, of 1 July 2020, 30 June 2020, and 28 June 2020. Finally, the problem formulation developed in this study is compared against prior work on an Earth-to-Mars interplanetary orbit transfer where it is found that the results of this research show significant improvement in transfer time relative to the prior work.
Ashley M. Korzun, Robert W. Maddock, Mark Schoenenberger, Karl T. Edquist, Carlie H. Zumwalt, Christopher D. Karlgaard
Journal of Spacecraft and Rockets pp 1-8; https://doi.org/10.2514/1.a35085

Abstract:
The Interior Exploration using Seismic Investigations, Geodesy and Heat Transport (InSight) mission touched down in Elysium Planitia on 26 November 2018, becoming NASA’s eighth successful entry, descent, and landing (EDL) at Mars. InSight inherited the successful 2008 Phoenix (PHX) EDL system, flying a nonspinning, ballistic trajectory with a 70 deg sphere–cone aeroshell (2.65 m diameter), a disk–gap–band parachute, and pulsed terminal descent and landing engines. InSight and Phoenix exhibited similar behavior, primarily in terms of trim attitude and an uninitiated roll reversal correlated with dynamic pressure, although the behaviors observed for InSight were more severe. For InSight, the initial clockwise roll and nonzero trim angle of attack were significant contributors to a short timeline, larger-than-predicted deceleration loads, and cross-track and up-track errors in landing location. The InSight and Phoenix reconstructions together indicate behavior more characteristic of a single, continuous instability region, suggesting that errors in the trim behavior from hypersonic nonequilibrium aerodynamics predictions along the dynamic pressure pulse may have a more substantial impact on nonspinning, ballistic entry vehicle performance.
Lorenzo Federici, Boris Benedikter,
Journal of Spacecraft and Rockets pp 1-12; https://doi.org/10.2514/1.a35076

Abstract:
This paper investigates the use of deep learning techniques for real-time optimal spacecraft guidance during terminal rendezvous maneuvers, in presence of both operational constraints and stochastic effects, such as an inaccurate knowledge of the initial spacecraft state and the presence of random in-flight disturbances. The performance of two well-studied deep learning methods, behavioral cloning (BC) and reinforcement learning (RL), is investigated on a linear multi-impulsive rendezvous mission. To this aim, a multilayer perceptron network, with custom architecture, is designed to map any observation of the actual spacecraft relative position and velocity to the propellant-optimal control action, which corresponds to a bounded-magnitude impulsive velocity variation. In the BC approach, the deep neural network is trained by supervised learning on a set of optimal trajectories, generated by routinely solving the deterministic optimal control problem via convex optimization, starting from scattered initial conditions. Conversely, in the RL approach, a state-of-the-art actor–critic algorithm, proximal policy optimization, is used for training the network through repeated interactions with the stochastic environment. Eventually, the robustness and propellant efficiency of the obtained closed-loop control policies are assessed and compared by means of a Monte Carlo analysis, carried out by considering different test cases with increasing levels of perturbations.
Joseph Hughes, Ryan Blay, Jack Ziegler, Phillip Anderson, Will Armijo, Jordan Maxwell, John Kendra
Journal of Spacecraft and Rockets pp 1-11; https://doi.org/10.2514/1.a35118

Abstract:
The Rotary-Motion Extended Array Synthesis spacecraft is a dual-spin spacecraft with long tethers held in tension by the spin. This paper derives the free-space (not considering orbital effects) dynamics of this spacecraft, and it compares two controllers in performance and the effort expended. In contrast to prior work, one of the controllers used takes account of the tether motion rather than just considering the hub motion. The spacecraft is simulated performing a challenging task including an acceleration and 90 deg slew with both controllers. The controller that accounts for the tether motion uses ∼10% less fuel while providing similar errors.
Alexander Puzev, Pini Gurfil, Vladimir Balabanov
Journal of Spacecraft and Rockets pp 1-12; https://doi.org/10.2514/1.a35042

Abstract:
CubeSats usually carry no propulsion system because of difficulties in scaling down existing propulsion technologies to meet the stringent size, volume, and power limits. The most challenging case in this context is the one-unit CubeSat: the use of which has been rapidly growing. To increase CubeSat functionality by allowing small orbit corrections, particularly due to orbit injection errors, this paper develops a new propulsion method relying on an ejection of masses from a reaction wheel. Using the reaction wheel’s kinetic energy, propellant masses are released from the wheel’s outer circumference using an impulsive application of an electric current. This paper models the dynamics of the satellite, the modified reaction wheel, and the ejected masses; evaluates the pros and cons of the proposed method; and provides a methodology for constructing a model of a mass-ejecting reaction wheel designed according to given specifications of a desired velocity change and satellite attitude stability. The system is useful for small orbit corrections, which can be used (for example) for CubeSat formation establishment after orbital injection.
Naoko Iwata, Sogo Nakanoya, Nobuyuki Nakamura, Noboru Takeda, Fumiya Tsutsui
Journal of Spacecraft and Rockets pp 1-11; https://doi.org/10.2514/1.a35030

Abstract:
This paper presents the experimental and analytical results of a heat rejection system consisting of a single-phase mechanically pumped loop and a space radiator. It has been developed for future crewed exploration missions to provide a large amount of heat dissipation capability. The radiator consists of a honeycomb radiator, eight aluminum pipes inserted through the honeycomb core, and two aluminum manifolds joined at both ends of the eight pipes. A graphite sheet is attached to the back of the aluminum skin, facing deep space to improve the thermal diffusivity in the in-plane direction. Silverized Teflon® tapes are put on the surface of the skin on the deep-space side, and the surface of the opposite skin is covered with multilayer insulation. A gap between the radiator surface and the flow channel protects against micrometeoroids and orbital debris. A thermal vacuum test was conducted to evaluate the thermal performance of the radiator. The coolant was HFE7200, and its flow rate was varied from about 50 to 200 kg/h. More than 50 temperatures across the radiator surface were measured for each case. Thermal Desktop® and SINDA/FLUINT software were used to develop the thermal mathematical model, which was correlated with the test results and evaluated through thermal analysis. The fin efficiency of the radiator increased with the mass flow rate of the coolant, reaching 0.9 or higher.
Shae T. Hart, , Peiman Naseradinmousavi
Journal of Spacecraft and Rockets pp 1-12; https://doi.org/10.2514/1.a35081

Abstract:
This paper presents a comparative study of six common pseudospectral (PS) methods for solving optimal spacecraft attitude control problems. The problems of minimum time, minimum control effort, and the linear combination thereof are considered here, for a rest-to-rest spacecraft slew maneuver. In this study, three traditional PS methods (Gauss, Legendre, and Chebyshev) and three variations of the novel Birkhoff PS method (Legendre–Gauss, Legendre–Gauss–Lobatto, and Chebyshev–Gauss–Lobatto) are used. The Wilcoxon rank sum test is used to perform statistical analysis to compare results between traditional and Birkhoff PS results. Accuracy, computational efficiency, and robustness to initial conditions are used as performance metrics for each PS method. The results show that for all three optimal control problems, the Birkhoff Legendre–Gauss method is more accurate than the traditional Gauss PS method. Furthermore, the results show that the Birkhoff variants could in some cases find the minimum cost using fewer iterations compared with traditional PS methods. The Birkhoff PS methods also are less susceptible to initial guesses for solutions that could cause the results to reach infeasible points or become trapped in local minimums.
Tristan Sarton du Jonchay, Hao Chen, Masafumi Isaji, Yuri Shimane, Koki Ho
Journal of Spacecraft and Rockets pp 1-16; https://doi.org/10.2514/1.a35094

Abstract:
This paper proposes an on-orbit servicing logistics optimization framework capable of performing the short-term operational scheduling and long-term strategic planning of sustainable servicing infrastructures that involve high-thrust, low-thrust, and/or multimodal servicers supported by orbital depots. The proposed framework generalizes the state-of-the-art on-orbit servicing logistics optimization method by incorporating user-defined trajectory models and optimizing the logistics operations with the propulsion technology and trajectory tradeoff in consideration. Mixed-integer linear programming is leveraged to find the optimal operations of the servicers over a given period, whereas the rolling horizon approach is used to consider a long time horizon accounting for the uncertainties in service demand. Several analyses are carried out to demonstrate the value of the proposed framework in automatically trading off the high- and low-thrust propulsion systems for both short-term operational scheduling and long-term strategic planning of on-orbit servicing infrastructures.
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