Shape memory alloys (SMAs) are unique materials that can be programmed to deform at a specific temperature and return to their original orientation. In this study, nitinol, a type of SMA, is used as an actuator for a deployable radiator on a CubeSat. The nitinol actuator is designed to unfold the radiator outward by 90 degrees. Key performance metrics-deployment angle, deployment time, and force generated-were measured to evaluate the system. These parameters were chosen to validate that the radiator deploys reliably, operates within energy constraints, and withstands accidental heating inside the spacecraft. The actuator consists of 14 nitinol joints, each 85 mm long, wrapped with 42 turns of nickel wire and insulated with Mylar to prevent heat loss. The 42 turns were selected to reach a resistance of 3 ohms using 0.72 amps and 2.12 volts. Each joint was tested with 1.5 watts of power to ensure simultaneous deployment. Two nitinols with different transformation temperatures were tested: the 60 degrees C actuator deployed in 10 minutes, while the 40 degrees C actuator deployed in 2 minutes. The 40 degrees C nitinol was chosen for final tests because of its faster response time, which better suits the CubeSat's need for quick deployment due to energy constraints.

To access the interstellar medium with current approaches requires 30 to 40 years, significantly longer than most mission lifetimes. The goal of this study was to explore mission concepts that will reach the interstellar medium in a primary mission’s lifetime (15 years or less). Faster access to the interstellar medium would allow high-capability science probes, with many relevant instruments, to explore the galaxy beyond our solar system in-situ. Science targets include the hydrogen wall structure, bow wave/shock, gravitational lens, foreground emissions and interstellar dust, just to name a few. Further, such a capability would enable rapid exploration of Kuiper belt objects in a much shorter time frame than current methods. Finally, distant targets include the solar gravitational lens, which may enable direct imaging of exoplanets.Figure 1: The Interstellar Medium Science TargetsWe examine a solar thermal propulsion (STP) system to rapidly access the local interstellar medium via a solar perihelion burn. This approach uses several Venus and Earth gravity assists to fly out to Jupiter and then would dive towards the Sun. Approaching within 3 solar radii a perihelion burn would be performed, maximizing the spacecraft’s ΔV to achieve high Solar System escape velocities. A unique aspect of the STP mission concept is that the Sun is not only used as a gravity well for an Oberth maneuver, but also to heat the fuel to ultra-high temperatures (>3000 K), enabling a monopropellant burn with high specific impulse (Isp). An in-depth modeling exercise found this approach to be preliminarily feasible, with escape velocities of around 9 AU/yr achievable with current technology, and up to 16 AU/yr with significant future technological advances.While the baseline STP design is capable of providing just under 9±1 AU/yr, Figure 2 highlights areas of key technological improvements that could be explored. Ultimately, if all technological paths could be implemented, the overall performance as a best-case scenario could reach approximately 16 AU/yr. Figure 2 also qualitatively ranks these improvements from most likely to least likely when reading the graph bottom up. For example, implementing turbopumps in the system is likely more readily feasible than reducing the liner thickness in the near future. It is assumed that these upgrades can be implemented in the future without incurring any additional mass penalty over the baseline design. Thus, it predicts best case performance, and actual values would likely be lower. The improvements could be the result of a single point improvement, or a propagation of several developments.Figure 2: Overview of Solar Thermal Propulsion PerformanceAfter reviewing the STP approach, and comparing it to a solid rocket motor (SRM), it was found that with currently available technology, SRM outperforms STP with an escape velocity of approximately 12 Au/yr. However, future advances in heat exchanger lining materials, turbo pumps, and advanced heat exchanger geometries may enable solar thermal propulsion to provide higher escape velocities, which would provide one of the fastest ways to exit the solar system. Of particular importance is heating the hydrogen to 3,500K. Using a perihelion burn as a kick stage for a nuclear electric propulsion system was found to be particularly effective for achieving even higher escape velocities, up to 19.5 Au/yr.

Oscillating heat pipes (OHPs) represent a promising advancement over traditional heat pipes, yet their operational boundaries, especially for long OHPs, remain insufficiently understood. This study investigates the impact of varying adiabatic length, channel diameter, and fill ratio on thermal performance, crucial for assessing their suitability for engineering applications like spacecraft thermal management. Three long OHPs, ranging from 451 mm to 770 mm in total length, were subjected to multiple performance tests, employing channel diameters of 1.1 mm and 1.9 mm, along with adiabatic lengths of 305 mm and 610 mm. The experimental setup involved mounting the OHPs onto a testbed, monitored by nine K-type thermocouples. The tests, conducted horizontally to eliminate gravity-assistance, revealed that thermal performance is significantly influenced by channel diameter, adiabatic length, and fill ratio. Notably, optimal performance was observed at a 50% fill ratio, while reductions in diameter hindered start-up at a 70% fill ratio and failed to start-up at 30% fill ratio. These findings highlight the limitations of long OHPs, which is crucial to determine the limits of their applicability and dimensional constraints.

Two-phase pumped cooling systems are applied when it is required to maintain a very stable temperature for heat dissipation in a system. A novel additively manufactured evaporator for two-phase thermal control was developed at NASA Jet Propulsion Laboratory (JPL). The Two-Phase Mechanically Pumped Loop (2PMPL) allows to manage the heat transfer with much wider breadth of control authority compared to capillary-based systems, while alleviating the system's sensitivity to pressure drops. The focus of this work is the understanding and capturing the micro-scale evaporation occurring in the porous structure of the evaporator. The Boiling and Phase Change Heat Transfer Laboratory at the University of California, Los Angeles (UCLA) developed an all-encompassing numerical simulation tool to predict the operational thermal behavior of the evaporator considering the effect of the liquid-vapor interface at the wick-to-vapor boundary. The numerical model incorporated the behaviour of the liquid-vapor meniscus at particle level located along the evaporative boundary between the wick structure and the vapor chamber. The numerical model allowed to study the effect of different parameters, such as boundary conditions, geometry, wick and fluid properties. An experimental setup was built at UCLA in order to characterize the heat transfer within an additively manufactured porous sample fabricated at JPL and in particular its evaporative heat load under certain heat inputs. The experimental efforts served as validation for the numerical results and aided in the characterization of the transient phenomena, such as dry-out.

Modeling thermal states for complex space missions, such as the surface exploration of airless bodies, requires high computation, whether used in ground-based analysis for spacecraft design or during onboard reasoning for autonomous operations. For example, a finite-element thermal model with hundreds of elements can take significant time to simulate, which makes it unsuitable for onboard reasoning during time-sensitive scenarios such as descent and landing, proximity operations, or in-space assembly. Further, the lack of fast and accurate thermal modeling drives thermal designs to be more conservative and leads to spacecraft with larger mass and higher power budgets. The emerging paradigm of physics-informed machine learning (PIML) presents a class of hybrid modeling architectures that address this challenge by combining simplified physics models with machine learning (ML) models resulting in models which maintain both interpretability and robustness. Such techniques enable designs with reduced mass and power through onboard thermal-state estimation and control and may lead to improved onboard handling of off-nominal states, including unplanned down-time. The PIML model or hybrid model presented here consists of a neural network which predicts reduced nodalizations (distribution and size of coarse mesh) given on-orbit thermal load conditions, and subsequently a (relatively coarse) finite-difference model operates on this mesh to predict thermal states. We compare the computational performance and accuracy of the hybrid model to a data-driven neural net model, and a high-fidelity finite-difference model of a prototype Earth-orbiting small spacecraft. The PIML based active nodalization approach provides significantly better generalization than the neural net model and coarse mesh model, while reducing computing cost by up to 1.7x compared to the high-fidelity model.

Future space exploration missions require advanced thermal control systems (TCS) to dissipate heat from spacecraft, rovers, or habitats to external environments. These TCSs must be lightweight, reliable, and able to effectively control cabin and equipment temperatures under widely varying heat loads and ambient temperatures. In contrast to single phase pumped coolant loops, two phase pumped loops are very attractive for this application because of the uniform cooling temperature provided by the boiling coolant, low required pumping power, high heat transfer coefficients, and high thermal conductance. However, introduction of two phase flow can pose design challenges associated with flow management and dynamic stability. A condensing radiator technology is needed to enable future heat rejection systems with high turndown ratio, compatibility with freezing, and deployability, while balancing considerations unique to two phase flow condensation. To meet these performance requirements, Creare has developed a freeze tolerant, variable conductance radiator for deployable heat rejection in two phase pumped loop systems.

High-power small satellites will play an important role in reducing the cost of space missions. High power levels are required to satisfy high bandwidth communication needs, as well as to accommodate other critical systems such as propulsion and high-power electronics. One of the current high-power limitations is the need to overcome the thermal challenges associated with high thermal loads. While substantial work has been done in the development of deployable solar arrays, relatively little attention has been given to small satellite deployable radiators. Previous deployable small satellite radiators rely on a mechanical hinge to conduct heat from spacecraft to radiator. However, this presents a thermal choke point and limits heat flow. Thus, a deployable radiator design concept is currently being explored as a thermal solution for high powered electronics in CubeSats. This Additively Manufactured Deployable Radiator Oscillating Heat Pipes (AMDROHP) design concept combines the function of a deployable radiator with high performance Oscillating Heat Pipes in this compact thermal solution. In this project, significant efforts have gone into developing the mechanical, deployable aspect of this design while maximizing its thermal performance, and this joint development has been discussed in previous publication. This initial design was additively manufactured and thermally tested at the Jet Propulsion Laboratory. The results of the thermal testing of the initial AMDROHP design are discussed and presented in this work. The device is tested across a range of heat inputs under “micro-gravity” and “gravity-assisted” orientations for the working fluids R134a and Ammonia. The performance and behavior of the AMDROHP device are characterized by transient temperature measurement data under these different conditions. The results were interpreted to determine the feasibility of the design. Although AMDROHP did operate under “gravity-assisted” orientation, it did not start-up under “micro-gravity” orientation. Furthermore, the range of operation under “gravity-assisted” orientation was less than expected. Based on these results, possible design changes have been identified to improve AMDROHP performance under space-like conditions. These changes include creating a shorter adiabatic length by decreasing the path length of the helical joint, as well as increasing the inner channel diameter. These changes will allow for better thermal performance and to better avoid any imperfections in the additive manufacturing process to cause negative effects on OHP operation. In this study, experimental testing provided actionable information about the initial design of AMDROHP to lead to design improvements. These design improvements will be implemented in the next design iteration of AMDROHP.

Natural pits on the Moon expose deep cross-sections of the lunar maria, enabling direct investigation of the Moon's volcanic history and providing potential access to subsurface lava tubes. The Moon Diver mission concept seeks to explore the Mare Tranquillitatis pit, which exposes the largest wall of bedrock of the mare pits (similar to 65 m). The concept is enabled by two innovative capabilities: pinpoint landing near the pit and robotic access to its nearvertical wall with an instrument package to examine the elemental chemistry, mineralogy, and morphology of these bedrock layers. Pinpoint landing uses closed-loop guidance with terrain-relative navigation (TRN), which was advanced by Perseverance landing on Mars, to deliver the lander within a 100-m ellipse. The Axel robotic explorer, which remains tethered to the lander, would egress onto the lunar surface and traverse the relatively flat terrain to the pit's funnel entrance. The lander, which is the data link to Earth, also serves as an anchor and provides power and communication to the rover through its tether. The rover is a novel two-wheeled platform with a trailing boom and a spool that pays out the tether as the rover traverses toward the pit. The 300-m long tether is well margined for the rover to scale the pit wall. The rover carries a surface preparation tool and three additional instrument types: (a) three high-resolution cameras for acquiring context images of the near and far walls with the near-wall pair in a stereoscopic configuration, (b) an alpha-particle-X-ray spectrometer (APXS) for elemental composition, and (c) a multi-spectral microscopic imager (MMI) that uses controlled lighting for mineralogy. The surface-preparation tool removes dust and patina from the rock wall by grinding a small area. This tool, the MMI, and the APXS would be deployed from an instrument bay inside the wheel wells. The rover would independently point each instrument at the same target on the wall with millimeter repeatability. Landing shortly after sunrise, the surface mission timeline is just shy of a lunar daytime (14 Earth days). Beyond the primary mission, the rover would be capable of descending from the overhang and peering into the void that may open to a large cave or lava tube. Lunar pits provide an exciting new target for exploration using innovative robotic capabilities that have been tested with integrated science instruments at multiple terrestrial analogue sites including a pit with basaltic layers in Arizona.

As small spacecraft technologies develop, thermal management devices need to meet the growing demands of high-powered electronics. Currently being developed to meet this demand in CubeSats is the Additively Manufactured Deployable Radiator Oscillating Heat Pipes (AMDROHP). AMDROHP seeks to implement the high thermal conductivity and two-phase technology of Oscillating Heat Pipes into a unique deployable radiator design for a 3U CubeSat, taking advantage of additive manufacturing capabilities. While much consideration has been put into designing the AMDROHP on its own as a heat exchanger, there is also the need for it to be evaluated thermally at a system-level with the rest of the CubeSat while in orbit. In this study, thermal orbital spacecraft simulations, through the Thermal Desktop software, were performed to analyze how AMDROHP thermally integrates and interacts with the rest of the CubeSat and evaluate the survivability of temperature-sensitive components on the spacecraft. The simulations in this study included an 11th-orbit beta angle sweep for a tumbling orientation of the spacecraft in Low Earth Orbit (LEO). These simulations were performed with two AMDROHP devices in the CubeSat bus, each under a direct 25W heat input and performing with a thermal conductance of 6 W/K, which corresponds to the projected performance of the AMDROHP device while in operation. In this paper, the Thermal Desktop model of the AMDROHP CubeSat includes all major physical components, connections, heat loads, and thermal and optical materials. Then, steps are taken to improve the computational speed of the model. Furthermore, the means of addressing the modeling of the complex two-phase behavior of the OHP is outlined. Then, a number of test cases considering various operating conditions were simulated. From these simulations, orbital temperatures of sensitive components, primarily electronics, were collected and analyzed to find the minimum and maximum operating temperatures across all potential orbits. These temperatures were then evaluated to determine the component’s survivability in a worst-case scenario in orbit. From the results, it was found that, with the projected conductance of AMDROHP, all components operate under safe temperature conditions for any beta angle while in Low Earth Orbit. The evaporator is consistently the hottest component of the spacecraft and electronics boards all maintain survivable temperatures and are not at risk of over or underheating, even at worst case temperatures for all orbits tested. Based on the results and analysis of this conceptual study, it is suggested that AMDROHP will perform as an effective management device for small satellites.

Oscillating Heat Pipes (OHPs) are unique two-phase heat transfer devices with many advantages over standard and more widely adopted thermal control devices. Specifically, OHPs are able to operate passively, over a wide temperature range, and under high heat fluxes, with few design constraints. OHPs operate based on the process of evaporating and condensing a working fluid on opposite ends of a series of serpentine channels. The pressure difference causes an oscillating behavior of the fluid to travel along the direction of the channels enabling a high heat transfer rate. The capability of this technology has garnered much attention and interest for a variety of applications in different fields. While as a thermal management device, OHP technology shows much promise, there is still much to be understood about the fundamental principles of its operation. One of the most important aspects of OHP operation that is still not well understood is the phenomena known as “start-up”. Start-up is when the proper conditions of an OHP are created such that the operating, oscillatory process of an OHP is enabled to begin. In this work, experimental testing is performed to investigate the relationship between evaporator and condenser length on OHP start-up. In this study, combinations of small, medium, and large length evaporators and condensers were used on a 42-turn OHP with 1 mm square channels, charged with R134a to 50% fill ratio with the goal to quantify the minimum heat that would initiate start-up. Before this test could be performed, issues in consistency of the start-up of OHP had to be resolved. A method to “reset” the liquid-vapor distribution to overcome any adverse history from previous OHP operation developed is discussed. Then, a zero-heat input method of confirming a liquid-vapor distribution favorable to start-up is outlined. After developing this method to achieve a consistent start-up heat input in the OHP, tests were performed for nine different configurations of large, medium, and small evaporator and condensers, to determine their relationship with the minimum heat required to start-up an OHP. It was observed that both the size of the evaporator and condenser both influence the start-up heat load. That is, a large condenser and evaporator can reduce the heat load required to start-up. Additionally, the size of a condenser is much more influential than the size of the evaporator. The overarching discovery is that the process of start-up relies heavily on the history of the OHP i.e., the initial liquid-vapor distribution of the working fluid in the channels. In measuring start-up, it is important to recognize and address how the current state of the OHP can lead to inconsistent results.

Two-phase mechanically pumped fluid loops have the potential to provide numerous benefits for spacecraft based applications. Here a hybrid two-phase capillary and mechanically pumped fluid loop that shows promise for spacecraft thermal control applications is analyzed and its operational limits are characterized. Data from an experimental system that incorporates a 3D printed evaporator is correlated to several models that together provide a comprehensive characterization of the system behavior. The loop performance is analyzed using pointwise as well volume-averaged governing equations and resistance network models, to predict flooding and dryout limits. Also, system-level numerical (including 3-D CFD simulations) transient and linear-response analyses are performed predicting the changing loop phase distributions and the observed time-dependent behavior. With the 3D-printed evaporator, a notably high evaporator conductance of 30 W/cm(2)-K and heat flux over 10 W/cm(2) over large area are achieved with negligible temperature non-uniformity. (C) 2021 Elsevier Ltd. All rights reserved.

This paper reports on experimental and analytical results showing the relationship between metal porous materials properties and laser parameters developed using a Powder Bed Fusion. We have 3D printed a series of porous materials coupons by varying laser parameters such as scan speed and laser power. This is shown to be a successful strategy for four different materials: AlSi10Mg, Stainless steel 316L, Inconel 625 and Ti-6Al-4V. The resulting porous material is characterized with measurements of the effective pore radius, permeability and the porosity to assess the performance. The experimental results lead to the definition of a process window of laser parameters that can be used to fabricate metal porous materials for each of the specified materials. Within this process window, we show that it is possible to obtain porous materials with pore radii ranging from 1.1 to 28.1 μm, porosity ranging from 10.4 to 42.4% and permeability ranging from 5.9 × 10–16 to 4.4 × 10–12 m2. This is solely achieved by varying laser parameters. We also describe models to predict porous material’s properties. Based on experimental results, we show that porosity can be controlled by varying laser parameters and, especially, applying appropriate energy density.

In this work, we explore two different Oscillating Heat Pipe (OHP) evaporator-condenser placement configurations, and investigate and quantify the effects on thermal performance. The proposed study focuses on two different OHP evaporator-condenser configurations with a change in adiabatic length. One of the challenges in OHP literature is the variety of experiment setups, (e.g. varying condenser, evaporator, and adiabatic lengths, heat input) which makes it difficult to compare results directly. To quantitatively compare thermal performance, a (or a set of) standardized metric(s) must be used. Therefore, we define a standardizing metric to quantify OHP’s ability to conduct heat that can be used across multiple experiments and setups. This study was conducted on an additively manufactured flat-plate AlSi10Mg OHP which has a channel diameter of 1.4 mm, 22 turns, and a plate size of 200 mm × 90 mm × 4 mm. Both the evaporator and condenser are rectangles with contact areas of 46 mm by 78 mm. The OHP was charged with R134a with a 45% filling ratio. In these tests, the location of the evaporator was fixed, while the placement of the condenser is varied such that the adiabatic length ranged between 4 mm to 94 mm. The condenser temperature was maintained between 10°C to 25°C and the heat input ranged between 20W to 50W. The results showed that a reduction in adiabatic length increased the thermal conductivity. To quantify the thermal performance, the thermal conductivities of an empty and charged OHP were determined for each placement configuration, then a thermal conductivity ratio of charged and empty OHP can be determined to quantify the improved performance. For an adiabatic length of 4 mm, we observed that the OHP’s ability to conduct heat was 40 times more effective when compared to an empty OHP. It was also observed that the OHP’s ability to conduct heat was 9 times more effective when compared to an empty OHP for an adiabatic length of 94 mm. We conclude that the area outside the evaporator-condenser that is neither heated nor cooled, called the reservoir, significantly influenced the thermal performance. The OHP with a shorter adiabatic length increased the reservoir in the condenser region which showed higher thermal performance. In this placement configuration, the reservoir essentially acted as an extension of the condenser. This is a favorable condition where the subcooled liquid slugs re-enter the condenser section which affects heat transfer drastically. Thus, the placement of the evaporator-condenser will influence OHP performance due to the reservoir and warrants future work.

The Far-side Array for Radio Science Investigations of the Dark ages and Exoplanets (FARSIDE) is a proposed mission concept to the lunar far side that seeks to deploy and operate an array of 128 dual-polarization, dipole antennas over a region of 100 square kilometers. The resulting interferometric radio telescope would provide unprecedented radio images of distant star systems, allowing for the investigation of faint ra-dio signatures of coronal mass ejections and energetic particle events and could also lead to the detection of magnetospheres around exoplanets within their parent star's habitable zone. Simultaneously, FARSIDE would also measure the ‘Dark Ages' of the early Universe at a global 21-cm signal across a range of red shifts (z ~50-100). Each discrete antenna node in the array is connected to a central hub (located at the lander) via a communication and power tether. Nodes are driven by cold-operable electronics that continuously monitor (for 5 or more years) an extremely wide-band of frequencies (200 kHz to 40 MHz), which surpass the capabilities of Earth-based telescopes by two orders of magnitude. Achieving this ground-breaking capability requires a robust deployment strategy on the lunar surface, which is feasible with existing, high-TRL technologies (demonstrated or under active development) and is capable of delivery to the surface on next-generation commercial landers, such as Blue Origin's Blue Moon Lander. This paper presents an antenna packaging, placement, and surface deployment trade study that leverages recent advances in tethered mobile robots under development at NASA's Jet Propulsion Laboratory, which are used to deploy a flat, antenna-embedded, tape tether with optical communication and power transmission capabilities. We investigate the feasibility of deploying 4 separate 12-km tethers, each with 64 remotely powered electronics nodes, using just four rovers that trace out a large spiral pattern that is over 10-km in diameter, where each arm in the spiral is precisely laid out and aligned with respect to a global heading (NSEW) to provide dual polarization across the entire array. Further, we provide a detailed design for an instrument/antenna-embedded tether, detail its accommodation within and deployment from a two-wheeled rover, and show how the entire system can be packaged and deployed from a commercial lunar lander system.

Europa is a premier target for advancing both planetary science and astrobiology, as well as for opening a new window into the burgeoning field of comparative oceanography. The potentially habitable subsurface ocean of Europa may harbor life, and the globally young and comparatively thin ice shell of Europa may contain biosignatures that are readily accessible to a surface lander. Europa's icy shell also offers the opportunity to study tectonics and geologic cycles across a range of mechanisms and compositions. Here we detail the goals and mission architecture of the Europa Lander mission concept, as developed from 2015 through 2020. The science was developed by the 2016 Europa Lander Science Definition Team (SDT), and the mission architecture was developed by the preproject engineering team, in close collaboration with the SDT. In 2017 and 2018, the mission concept passed its mission concept review and delta-mission concept review, respectively. Since that time, the preproject has been advancing the technologies, and developing the hardware and software, needed to retire risks associated with technology, science, cost, and schedule.

The harsh environment of Venus poses a significant challenge for the design of lightweight surface probes. The authors investigated isogrid-stiffened spherical shells produced by additive manufacturing. A Ti-6Al-4V, 1.12-m-diam lightweight structural shell capable of surviving a Venus surface environment was designed and assessed via analytical methods and finite element modeling. Subscale components (200 mm diameter) were designed, analyzed, and fabricated using laser powder bed fusion and tested under hydrostatic external pressure at ambient temperatures for model validation. A subscale sphere was successfully tested under Venus-like pressures and temperatures at a hot isostatic press facility. It was found that additively manufactured isogrid shells could lead to significant mass, cost, and schedule savings, freeing resources for increased science return from future missions.

The interstellar medium (ISM) represents the next frontier in space exploration, with many new discoveries to be made. The challenge, being so far away from Earth, the ISM requires many decades to reach. To advance our knowledge of what exists beyond our solar system, new approaches for rapid access are required. One such approach is solar thermal propulsion (STP). The approach uses several Venus and Earth gravity assists to fly to Jupiter and use its gravity well to dive towards the Sun. Approaching within three solar radii a perihelion burn would be performed, maximising the spacecraft’s ΔV to achieve high solar system escape velocities. A unique aspect of the STP mission concept is that the Sun is used not only as a gravity well for an Oberth manoeuvre, but also to heat the fuel to ultra-high temperatures (> 3000 K), enabling a monopropellant burn with high specific impulse (Isp). Prior preliminary studies indicated escape velocities of over 20 astronomical unit (AU)/year would be possible. An in-depth modelling exercise was undertaken to determine how such a system would perform. The model in this paper showed the current STP design is capable of providing just under 9 ± 1 AU/year, but there are many technology developments that could increase escape velocity. The technologies vary from items that could be implemented in the near term, like turbo-pumps driven by the hydrogen, to items requiring more extensive development programs like thin coatings which do not erode in superheated hydrogen. After reviewing the STP approach, and comparing it to a solid rocket motor (SRM), it was found that with currently available technology, SRM outperforms STP with an escape velocity of approximately 10–12 AU/year. However, future advances in heat exchanger lining materials, turbo pumps, and advanced heat exchanger geometries may enable solar thermal propulsion to provide higher escape velocities, providing one of the fastest ways to exit the solar system. Ultimately, if all technology paths could be implemented with minimal side effects, the performance in a best-case scenario could reach up to 16 AU/year.

Metal materials are used widely in industry for piping, as they provide desirable material properties for use in harsh environments. In this work, we explore different flexible pipe joint designs that enable fluid transport in flexible heat transfer devices. Flexible joints would allow heat transfer devices to fit in variable cooling areas. The proposed flexible joints will be 3D printed in metal with the goal of maximizing their bending angles, while minimizing volume. The objective of this research is to develop flexible pipes to connect parts involving fluid flow, and we propose three types of flexible joint designs: 1) helical coil spring, 2) torsional coil spring, and 3) serpentine structure. Then, we introduce a low-cost design process to improve these joint designs. Metal 3D printing offers increased design flexibility in comparison to traditional metal manufacturing methods. However, metal 3D printing can be labor intensive and costly, thus the number of design iterations using metal printing should be minimized during the design process. To reduce the frequency of metal prototyping, we used plastic 3D prints for rapid prototyping and evaluation. The flexible joint design considered the effects of pipe shape, pipe size, pipe diameter, free length, and total pipe length. Although mathematical models exist for the three types of flexible designs, they tend to be complex, difficult to implement, and specific to certain types of boundary conditions. In our experience, it is faster and more meaningful to perform experimental parametric studies using plastic prints due to accessibility and cost-effectiveness of plastic 3D printers. However, analytical models do offer significant insights during the design process. Thus, we developed simple analytical models for each of the three types of flexible joint designs which can be used in the initial design phase. Tests were performed to characterize the designs. Then, finite element simulations were performed for plastic prints and these simulations were validated against experimental data. We show that the results from analytical models, finite element simulations, and testing results for plastic prints are consistent, and these tools can be used to predict performance. In this paper, we discuss the following findings: 1) development of simple analytical models, including successes, limitations, and challenges; 2) the role of finite element simulations in the design process; 3) testing results from the three types of joint design; 4) quantification, interpretation, and discussion of testing results.

This paper presents a three-dimensional heat transfer analysis of a flat-plate oscillating heat pipe (OHP). To enhance implementation of OHPs in various applications, it is important to understand the comprehensive phenomena that include thermal diffusion in the whole structure and thermal hydraulics in the channel. We developed an OHP model that includes the effect of thermal diffusion and thermo-fluid behavior. The model was validated with experimental results of a flat-plate aluminum OHP which has a size of 200 mm x 90 mm x 3.8 mm, a channel diameter of 1.0 mm, and a turn number of 42. The effect of surface roughness in the channel and liquid film thickness on operating temperature is investigated. The relation between thermo-fluid behavior in the channel and temperature distribution of the OHP is analyzed. For aluminum OHP, the temperature difference across the thickness direction is 0.1 ?C which is relatively small, whereas, the moving hot spot: local hightemperature region, is found in the planer surface. The maximum hot spot superheat temperature reached 5.4 ?C for 200 W. The novelties of this paper are to develop the three-dimensional comprehensive OHP model and to reveal the combined effect of thermo-hydraulic phenomena and thermal diffusion in OHP structure.