A combined experimental and numerical study of multidimensional viscous flows at speeds exceeding 8 km/s is reported. Experiments were performed in the X3 superorbital expansion tube with air and nitrogen test flows at a Mach number and total enthalpy of 10 and 40 MJ/kg, respectively. Laminar skin friction, heat flux and pressure measurements were obtained at regular intervals along one wall of a rectangular duct. The spatial resolution of the transducers was chosen to capture the multidimensional flow phenomena within the duct. Quasi-steady flow periods were established along the entire length of the duct in the test times offered by the expansion tube. Direct skin friction measurements were accomplished through the use of ‘in house’ acceleration compensated transducers. The successful operation of these skin friction transducers in a high performance expansion tube was demonstrated. Furthermore, the systematic uncertainty in measured shear stress was significantly reduced with the development of a new pressure calibration technique. For the conditions tested, Reynolds analogy was shown to be valid to within experimental uncertainty. The experimental data was in excellent agreement with numerical estimates. Three-dimensional numerical simulations of the diverging duct revealed that the flowfield structure in the vicinity of the corners differs from that of an unbounded corner or a constant area duct. Real gas effects other than those present in the residual nonequilibrium levels of freestream dissociation were negligible for the conditions tested. A computational study of two waverider configurations recently tested in the X3 superorbital expansion tube was conducted to assist in the interpretation of past results. The off-design aerodynamic performance was also analyzed and showed that blunting the leading edges dramatically degraded the performance by increasing drag and decreasing lift for the conditions considered.

A combined experimental and numerical study of multidimensional viscous flows at speeds exceeding 8 km/s is reported. Experiments were performed in the X3 superorbital expansion tube with air and nitrogen test flows at a Mach number and total enthalpy of 10 and 40 MJ/kg, respectively. Laminar skin friction, heat flux and pressure measurements were obtained at regular intervals along one wall of a rectangular duct. The spatial resolution of the transducers was chosen to capture the multidimensional flow phenomena within the duct. Quasi-steady flow periods were established along the entire length of the duct in the test times offered by the expansion tube. Direct skin friction measurements were accomplished through the use of ‘in house’ acceleration compensated transducers. The successful operation of these skin friction transducers in a high performance expansion tube was demonstrated. Furthermore, the systematic uncertainty in measured shear stress was significantly reduced with the development of a new pressure calibration technique. For the conditions tested, Reynolds analogy was shown to be valid to within experimental uncertainty. The experimental data was in excellent agreement with numerical estimates. Three-dimensional numerical simulations of the diverging duct revealed that the flowfield structure in the vicinity of the corners differs from that of an unbounded corner or a constant area duct. Real gas effects other than those present in the residual nonequilibrium levels of freestream dissociation were negligible for the conditions tested. A computational study of two waverider configurations recently tested in the X3 superorbital expansion tube was conducted to assist in the interpretation of past results. The off-design aerodynamic performance was also analyzed and showed that blunting the leading edges dramatically degraded the performance by increasing drag and decreasing lift for the conditions considered.

<p> Orbital Debris is a growing challenge for the Space Industry. The increasing density of derelict objects in high-value orbital regimes is resulting in more conjunction warnings and break-up events with cascading repercussions on active satellites and spacecraft. The recent rapid growth of the commercial space industry, in particular proliferated satellite constellations, has placed orbital debris remediation at the forefront of Space Industry efforts. The need to remove existing debris, combined with a growing demand for active satellite life extension services, has created an emerging market for space logistics, in particular spacecraft capable of rendezvous and docking, orbital refueling, debris deorbiting, or object relocation. This market has seen numerous companies emerge with multi-purpose on-orbit servicing platforms. This ecosystem poses technological, economical, and policy questions to decision-makers looking to acquire platforms or invest in technologies and requires a System-of-Systems approach to determine mission and system concepts of merit. An open-source modeling, analysis, and simulation software toolbox has been developed which enables rapid early-stage analysis and design of diverse fleets of on-orbit servicing platforms, with a specific emphasis on active debris removal applications. The toolbox provides fetching and processing of real-time orbital catalog data, clustering and scoring of high-value debris targets, flexible and efficient multi-vehicle multi-objective time-varying routing optimization, and fleet-level lifecycle cost estimation. The toolbox is applied to a diverse sample of promising commercial platforms to enable government decision-makers to make sound investment and acquisition decisions to support the development of ADR technologies, missions, and companies. </p>

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<p>Multiple Signals of Opportunity (multi-SoOp) across different frequencies and polarizations</p> <p>offer a potential breakthrough for remote sensing of root-zone soil moisture (RZSM). Deeper penetration depths of existing communication transmissions in the frequency ranges of 137–138, 240–270, and 360–380 MHz enable the estimation of RZSM by complementing global navigation satellite system reflectometry (GNSS-R) in L-band. The small form factor of the multi-SoOp observatory allows for high spatiotemporal coverage of RSZM by a satellite constellation in a cost-effective manner. This study aims to develop models and tools to define mission requirements for various system parameters that affect observation accuracy and coverage, for the advancement of spaceborne multi-SoOp remote sensing. These parameters include frequency and polarization combinations, observation error, inter-frequency temporal coincidence, and configuration of the satellite constellation. We present the development of a retrieval algorithm and the sensitivity analysis of retrieval accuracy. The retrieval algorithm was evaluated using synthetic observations generated from multiyear time series of in-situ soil moisture (SM) and satellite-based vegetation data. The combined use of both high and low frequencies improves retrieval accuracy by limiting uncertainties from vegetation and surface SM and providing sensitivity to deeper layers. A bivariate model, derived from the sensitivity analysis, facilitates error prediction for future science missions. We introduce a framework for tradespace exploration of the multi-SoOp satellite constellation. A constellation design study indicates that a Walker constellation comprising 24 satellites with 3 orbital planes at 500 km and 50° inclination optimizes the coverage and mission cost under mission requirements. A tower-based field experiment validated the performance of a prototype antenna for multi-SoOp using the interference pattern technique. More field experiments with improved instruments are required to further advance the multi-SoOp technique.</p>

<p dir="ltr">In the pursuit of establishing a sustainable space economy within the cislunar region, it is vital to formulate transfer design strategies that uncover economically viable highways between different regions of the space domain. The inherent complexity of spacecraft dynamics in the cislunar space poses challenges in determining feasible transfer options. However, the motion characterized by known dynamical structures modeled through the circular restricted three-body problem (CR3BP) aids in the identification of pathways with reasonable maneuver costs and flight times. A framework is proposed that incorporates a quasi-periodic orbit (QPOs) as an option to design transfer scenarios. This investigation focuses on the construction of transfers between periodic orbits. The framework is exemplified by the construction of pathways between an L2 9:2 synodic resonant Near-Rectilinear Halo Orbit (NRHO) and a planar Moon-centered Distant Retrograde Orbit (DRO). The innate difference in the geometries of the departure and arrival orbits of the sample case, along with the lack of natural flows towards and away from them, imply that links between these orbits may necessitate costly maneuvers. A strategy is formulated that leverages the stable and unstable manifolds associated with intermediate periodic orbits and quasi-periodic orbits to construct end-toend trajectories. As part of this strategy, a systematic methodology is outlined to streamline the determination of transfer options provided by the 5-dimensional manifolds associated with a QPO family. This approach reveals multiple local basins of solutions, both interior and exterior-types, characterized by selected intermediate orbits. The construction of transfers informed by the manifolds associated with QPOs is more intricate than those based on periodic orbits. However, QPO-derived solutions allow for the recognition of alternative local basins of solutions and often offer more cost-effective transfer options when compared to trajectories designed using periodic orbits that underlie the QPOs.</p>

<p>  </p> <p>  </p> <p>An approach to modify surface flashover of insulators in vacuum by limiting duration of its high-current stage responsible for the damaging effects of a classic flashover was developed. The flashover assembly was made by TorrSeal-gluing copper electrodes (10 x 10 x 0.5 mm) to both side of an alumina ceramic sheet (0.635 mm thick). The modified flashover, referred to as low energy surface flashover (LESF), was achieved by utilization of a high voltage (HV) nanosecond pulser or addition of a resistor in series with the LESF assembly when HV DC was utilized. The duration of LESF was visualized by ICCD fast photography to be 100 – 200 ns accompanying electrical characteristics measurements, which gave insight of a way to control the flashover duration by inserting additional capacitor in parallel with the LESF assembly to increase the stored energy prior to breakdown. The LESF assembly was tested for > 1.5 million consecutive pulses and remained operational, while operation in high energy regime with parallel capacitor (4nF) lead to significant damage after 200 pulses.</p> <p>The igniting capabilities of LESF assembly was demonstrated via successful triggering of vacuum arc and a prototype pulsed plasma accelerator. The plasma plume propagation speed and angular distribution was measured via Langmuir probes. Efforts were made for temporally resolved spectroscopy measurements. </p> <p>The LESF assembly was improved by replacing TorrSeal-gluing with direct bonding of copper to alumina ceramic and changing the configuration from parallel plate to coaxial. The improved assembly was demonstrated to be operational throughout and after an extended test of 10 million pulses. A higher resolution ICCD photography revealed finer LESF discharge features including initial bright line across the insulator developing into a double-jet plasma plume propagating at around 10⁵m/s and later-on point-like attachment of the discharge column to the electrodes. The composition of the plasma and erosion pattern on the LESF assembly was studied via SEM/EDX analysis, which supported the predominant ceramic erosion over copper electrodes erosion.</p>

Aerocapture offers a near propellantless and quick method of orbit insertion at atmosphere bearing planetary destinations. Compared to conventional propulsive insertion, the primary advantage of using aerocapture is the savings in propellant mass which could be used to accommodate more useful payload. To protect the spacecraft from the aerodynamic heating during the maneuver, the spacecraft must be enclosed in a protective aeroshell or deployable drag device which also provides aerodynamic control authority to target the desired conditions at atmospheric exit. For inner planets such as Mars and Venus, aerocapture offers a very attractive option for inserting small satellites or constellations into very low circular orbits such as those used for imaging or radar observations. The large amount of propellant required for orbit insertion at outer planets such as Uranus and Neptune severely limits the useful payload mass that can delivered to orbit as well as the achievable flight time. For outer planet missions, aerocapture opens up an entirely new class of short time of flight trajectories which are infeasible with propulsive insertion. A systems framework for rapid conceptual design of aerocapture missions considering the interdependencies between various elements such as interplanetary trajectory and vehicle control performance for aerocapture is presented. The framework provides a step-by-step procedure to formulate an aerocapture mission starting from a set of mission objectives. At the core of the framework is the ``aerocapture feasibility chart", a graphical method to visualize the various constraints arising from control authority requirement, peak deceleration, stagnation-point peak heat rate, and total heat load as a function of vehicle aerodynamic performance and interplanetary arrival conditions. Aerocapture feasibility charts have been compiled for all atmosphere-bearing Solar System destinations for both lift and drag modulation control techniques. The framework is illustrated by its application to conceptual design of a Venus small satellite mission and a Flagship-class Neptune mission using heritage blunt-body aeroshells. The framework is implemented in the Aerocapture Mission Analysis Tool (AMAT), a free and open-source Python package, to enable scientists and mission designers perform rapid conceptual design of aerocapture missions. AMAT can also be used for rapid Entry, Descent, and Landing (EDL) studies for atmospheric probes and landers at any atmosphere-bearing destination.

<p>As the small satellite market has grown from a niche of the space economy to a full commercial force,  microthrusters remain an area of significant growth in the space industry as new technologies mature. The \textit{Film-Evaporation Microelectricalmechanical Tunable Array} (FEMTA) is one such device. FEMTA is \textit{microelectricalmechanical system} (MEMS) device that harnesses the microcapillary action of water and vacuum boiling to generate thrust. The water propellant is not chemically altered at all by the process; it is simply evaporated. This technology has been tested in relevant laboratory environments, and a suborbital flight opportunity in 2023 as a payload on a Blue Origin New Shepard rocket  will grant FEMTA a demonstration in a space environment. The flight will provide 150 seconds of weightlessness at the zenith of the suborbital flight path before the booster returns to land. During weightlessness, the experiment will be exposed to the ambient environment allowing for a full capability test of the thruster. The experiment is meant to demonstrate the propellant management system for FEMTA in 0G and measure the thrust produced by a FEMTA thruster.</p> <p><br></p> <p>The propellant management system portion of the experiment consists of an oversized version of the subsystem intended for use in the thruster. The propellant management system uses a hydrofluoroether to inflate a diaphragm to ensure constant wetting of the propellant tank exit and nozzle inlet. The experiment will take tank pressure data and flow sensor data to understand the system's behavior. The system is duplicated for redundancy and to double the possible data. This system requires further testing before being prepared for launch, vibrational testing, thermal testing, and vacuum testing. </p> <p><br></p> <p>The 0G thrust experiment and plume analysis portion of the experiment consists of numerical modeling and a novel thrust measurement approach. \textit{Direct Simulation Monte Carlo} (DSMC) is being applied to understand the pressure, density, and temperature distributions of the plume of water vapor produced by the FEMTA thruster. The FEMTA nozzle environment is challenging to simulate with computational fluid dynamics  or DSMC due to chaotic transient effects and because both the continuum and molecular regimes must be considered. The current analysis consisted of a two-dimensional model and investigated the effect of meniscus location and contact angle on thrust generated.</p> <p><br></p> <p>It is not possible to use traditional thrust measurement devices (sensitive torsional thrust stands or microsensors intended for use on small satellites) for microthrusters on a rocket booster. Two  novel approaches for performing thrust measurement in the range of 100 microNewtons have been investigated. The first approach ionizes the FEMTA thruster plume and analyzes the plasma by optical emission spectroscopy. The theory states that the relative intensity of a given wavelength observed correlates to the density of the species in the plasma. The density of water would be directly correlated to the thrust generated by FEMTA during the experiment, as more water is evaporated as thrust is increased. This method is no longer being considered for the suborbital experiment but did yield promising results. </p> <p><br></p> <p>The second approach employs a d'Arsonval meter, a photo-interrupt, and an Arduino controller. The d'Arsonval meter consists of a stationary permanent magnet with a moving coil and a pointer. Increasing the voltage in the coil causes a torque on the system due to the magnetic field induced by the permanent magnet. This torque causes a deflection of the pointer that is proportional to the voltage applied. The flag of the sensor will be placed in the path of the gas jet from the thruster. The force caused by the jet pressure will move the flag. An Arduino controller will vary the voltage to hold the flag in place. As the mass flow rate increases, the reaction force required to hold the flag in place will increase. This sensor can be calibrated using an analog cold gas system that passes various gases (air nitrogen, argon, etc.) through an orifice nozzle at mass flow rates that are set by a mass flow rate controller. DSMC analysis has been performed to understand the flow field and flow properties and how they directly relate to the force experienced by the flag sensor. </p> <p>An undergraduate course has supported parts of the work described in this dissertation. This course has applied the Vertically Integrated Projects approach to project-based learning. This method and its results were analyzed and lessons learned as well as a blueprint for future application of this method to other small satellite projects are discussed.</p>

<div>The orbits of interest for potential missions are stable or nearly stable to maintain long term presence for conducting scientific studies and to reduce the possibility of rapid departure. Near Rectilinear Halo Orbits (NRHOs) offer such stable or nearly stable orbits that are defined as part of the L1 and L2 halo orbit families in the circular restricted three-body problem. Within the Earth-Moon regime, the L1 and L2 NRHOs are proposed as long horizon trajectories for cislunar exploration missions, including NASA's upcoming Gateway mission. These stable or nearly stable orbits do not possess well-distinguished unstable and stable manifold structures. As a consequence, existing tools for stationkeeping and transfer trajectory design that exploit such underlying manifold structures are not reliable for orbits that are linearly stable. The current investigation focuses on leveraging stretching direction as an alternative for visualizing the flow of perturbations in the neighborhood of a reference trajectory. The information supplemented by the stretching directions are utilized to investigate the impact of maneuvers for two contrasting applications; the stationkeeping problem, where the goal is to maintain a spacecraft near a reference trajectory for a long period of time, and the transfer trajectory design application, where rapid departure and/or insertion is of concern.</div><div><br></div><div>Particularly, for the stationkeeping problem, a spacecraft incurs continuous deviations due to unmodeled forces and orbit determination errors in the complex multi-body dynamical regime. The flow dynamics in the region, using stretching directions, are utilized to identify appropriate maneuver and target locations to support a long lasting presence for the spacecraft near the desired path. The investigation reflects the impact of various factors on maneuver cost and boundedness. For orbits that are particularly sensitive to epoch time and possess distinct characteristics in the higher-fidelity ephemeris model compared to their CR3BP counterpart, an additional feedback control is applied for appropriate phasing. The effect of constraining maneuvers in a particular direction is also investigated for the 9:2 synodic resonant southern L2 NRHO, the current baseline for the Gateway mission. The stationkeeping strategy is applied to a range of L1 and L2 NRHOs, and validated in the higher-fidelity ephemeris model.</div><div><br></div><div>For missions with potential human presence, a rapid transfer between orbits of interest is a priority. The magnitude of the state variations along the maximum stretching direction is expected to grow rapidly and, therefore, offers information to depart from the orbit. Similarly, the maximum stretching in reverse time, enables arrival with a minimal maneuver magnitude. The impact of maneuvers in such sensitive directions is investigated. Further, enabling transfer design options to connect between two stable orbits. The transfer design strategy developed in this investigation is not restricted to a particular orbit but applicable to a broad range of stable and nearly stable orbits in the cislunar space, including the Distant Retrograde Orbit (DROs) and the Low Lunar Orbits (LLO) that are considered for potential missions. Examples for transfers linking a southern and a northern NRHO, a southern NRHO to a planar DRO, and a southern NRHO to a planar LLO are demonstrated.</div>