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Search indexed NASA NTRS and DOE OSTI research on propulsion, heat transfer, battery materials and energy systems. Follow report and document links to the original sources.

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NASA NTRS2011

Low Cost Nuclear Thermal Rocket Cermet Fuel Element Environment Testing

Deep space missions with large payloads require high specific impulse and relatively high thrust to achieve mission goals in reasonable time frames.1,2 Conventional storable propellants produce average specific impulse. Nuclear thermal rockets capable of producing high specific impulse are proposed. Nuclear thermal rockets employ heat produced by fission reaction to heat and therefore accelerate hydrogen, which is then forced through a rocket nozzle providing thrust. Fuel element temperatures are very high (up to 3000 K), and hydrogen is highly reactive with most materials at high temperatures. Data covering the effects of high-temperature hydrogen exposure on fuel elements are limited.3 The primary concern is the mechanical failure of fuel elements that employ high-melting-point metals, ceramics, or a combination (cermet) as a structural matrix into which the nuclear fuel is distributed. The purpose of the testing is to obtain data to assess the properties of the non-nuclear support materials, as-fabricated, and determine their ability to survive and maintain thermal performance in a prototypical NTR reactor environment of exposure to hydrogen at very high temperatures. The fission process of the planned fissile material and the resulting heating performance is well known and does not therefore require that active fissile material be integrated in this testing. A small-scale test bed designed to heat fuel element samples via non-contact radio frequency heating and expose samples to hydrogen is being developed to assist in optimal material and manufacturing process selection without employing fissile material. This paper details the test bed design and results of testing conducted to date.

Bradley, D. E.↗
NASA NTRS2005

Empirical Scaling Laws of Rocket Exhaust Cratering

When launching or landing a space craft on the regolith of a terrestrial surface, special attention needs to be paid to the rocket exhaust cratering effects. If the effects are not controlled, the rocket cratering could damage the spacecraft or other surrounding hardware. The cratering effects of a rocket landing on a planet's surface are not understood well, especially for the lunar case with the plume expanding in vacuum. As a result, the blast effects cannot be estimated sufficiently using analytical theories. It is necessary to develop physics-based simulation tools in order to calculate mission-essential parameters. In this work we test out the scaling laws of the physics in regard to growth rate of the crater depth. This will provide the physical insight necessary to begin the physics-based modeling.

Donahue, Carly M.↗
NASA NTRS2012

A Brief Historical Survey of Rocket Testing Induced Acoustic Environments at NASA SSC

A survey was conducted of all the various rocket test programs that have been performed since the establishment of NASA Stennis Space Center. The relevant information from each of these programs were compiled and used to quantify the theoretical noise source levels using the NASA approved methodology for computing "acoustic loads generated by a propulsion system" (NASA SP ]8072). This methodology, which is outlined in Reference 1, has been verified as a reliable means of determining the noise source characteristics of rocket engines. This information is being provided to establish reference environments for new government/business residents to ascertain whether or not their activities will generate acoustic environments that are more "encroaching" in the NASA Fee Area. In this report, the designation of sound power level refers to the acoustic power of the rocket engine at the engine itself. This is in contrast to the sound pressure level associated with the propagation of the acoustic energy in the surrounding air. The first part of the survey documents the "at source" sound power levels and their dominant frequency bands for the range of engines tested at Stennis. The second part of the survey discusses how the acoustic energy levels will propagate non ]uniformly from the test stands. To demonstrate this, representative acoustic sound pressure mappings in the NASA Stennis Fee Area were computed for typical engine tests on the B ]1 and E ]1 test stands.

Allgood, Daniel C.↗
NASA NTRS2012

Probable Rotation States of Rocket Bodies in Low Earth Orbit

In order for Active Debris Removal to be accomplished, it is critically important to understand the probable rotation states of orbiting, spent rocket bodies. As compared to the question of characterizing small unresolved debris, in this problem there are several advantages: (1) objects are of known size, mass, shape and color, (2) they have typically been in orbit for a known period of time, (3) they are large enough that resolved images may be obtainable for verification of predicted orientation, and (4) the dynamical problem is simplified to first order by largely cylindrical symmetry. It is also nearly certain for realistic rocket bodies that internal friction is appreciable in the case where residual liquid or, to a lesser degree, unconsolidated solid fuels exist. Equations of motion have been developed for this problem in which internal friction as well as torques due to solar radiation, magnetic induction, and gravitational gradient are included. In the case of pure cylindrical symmetry, the results are compared to analytical predictions patterned after the standard approach for analysis of symmetrical tops. This is possible because solar radiation and gravitational torques may be treated as conservative. Agreement between results of both methods ensures their mutual validity. For monotone symmetric cylinders, solar radiation torque vanishes if the center of mass resides at the geometric center of the object. Results indicate that in the absence of solar radiation effects, rotation states tend toward an equilibrium configuration in which rotation is about the axis of maximum inertia, with the axis of minimum inertia directed toward the center of the earth. Solar radiation torque introduces a modification to this orientation. The equilibrium state is asymptotically approached within a characteristic timescale given by a simple ratio of relevant characterizing parameters for the body in question. Light curves are simulated for the expected asymptotic final rotation states of model objects, and these are compared to data derived from physical models of the same objects, tested in the Optical Measurements Center at JSC. Comparison to relevant light curves from actual orbiting rocket bodies are also performed, and diagnostic features of such curves are examined.

Ojakangas, Gregory W.↗
NASA NTRS2012

Mechanical and Combustion Performance of Multi-Walled Carbon Nanotubes as an Additive to Paraffin-Based Solid Fuels for Hybrid Rockets

Paraffin-based solid fuels for hybrid rocket motor applications are recognized as a fastburning alternative to other fuel binders such as HTPB, but efforts to further improve the burning rate and mechanical properties of paraffin are still necessary. One approach that is considered in this study is to use multi-walled carbon nanotubes (MWNT) as an additive to paraffin wax. Carbon nanotubes provide increased electrical and thermal conductivity to the solid-fuel grains to which they are added, which can improve the mass burning rate. Furthermore, the addition of ultra-fine aluminum particles to the paraffin/MWNT fuel grains can enhance regression rate of the solid fuel and the density impulse of the hybrid rocket. The multi-walled carbon nanotubes also present the possibility of greatly improving the mechanical properties (e.g., tensile strength) of the paraffin-based solid-fuel grains. For casting these solid-fuel grains, various percentages of MWNT and aluminum particles will be added to the paraffin wax. Previous work has been published about the dispersion and mixing of carbon nanotubes.1 Another manufacturing method has been used for mixing the MWNT with a phenolic resin for ablative applications, and the manufacturing and mixing processes are well-documented in the literature.2 The cost of MWNT is a small fraction of single-walled nanotubes. This is a scale-up advantage as future applications and projects will require low cost additives to maintain cost effectiveness. Testing of the solid-fuel grains will be conducted in several steps. Dog bone samples will be cast and prepared for tensile testing. The fuel samples will also be analyzed using thermogravimetric analysis and a high-resolution scanning electron microscope (SEM). The SEM will allow for examination of the solid fuel grain for uniformity and consistency. The paraffin-based fuel grains will also be tested using two hybrid rocket test motors located at the Pennsylvania State University s High Pressure Combustion Lab.

Larson, Daniel B.↗
NASA NTRS2012

Formulation, Casting, and Evaluation of Paraffin-Based Solid Fuels Containing Energetic and Novel Additives for Hybrid Rockets

This investigation studied the inclusion of various additives to paraffin wax for use in a hybrid rocket motor. Some of the paraffin-based fuels were doped with various percentages of LiAlH4 (up to 10%). Addition of LiAlH4 at 10% was found to increase regression rates between 7 - 10% over baseline paraffin through tests in a gaseous oxygen hybrid rocket motor. Mass burn rates for paraffin grains with 10% LiAlH4 were also higher than those of the baseline paraffin. RDX was also cast into a paraffin sample via a novel casting process which involved dissolving RDX into dimethylformamide (DMF) solvent and then drawing a vacuum on the mixture of paraffin and RDX/DMF in order to evaporate out the DMF. It was found that although all DMF was removed, the process was not conducive to generating small RDX particles. The slow boiling generated an inhomogeneous mixture of paraffin and RDX. It is likely that superheating the DMF to cause rapid boiling would likely reduce RDX particle sizes. In addition to paraffin/LiAlH4 grains, multi-walled carbon nanotubes (MWNT) were cast in paraffin for testing in a hybrid rocket motor, and assorted samples containing a range of MWNT percentages in paraffin were imaged using SEM. The fuel samples showed good distribution of MWNT in the paraffin matrix, but the MWNT were often agglomerated, indicating that a change to the sonication and mixing processes were required to achieve better uniformity and debundled MWNT. Fuel grains with MWNT fuel grains had slightly lower regression rate, likely due to the increased thermal conductivity to the fuel subsurface, reducing the burning surface temperature.

Larson, Daniel B.↗
NASA NTRS2012

Hydrocarbon Rocket Technology Impact Forecasting

Ever since the Apollo program ended, the development of launch propulsion systems in the US has fallen drastically, with only two new booster engine developments, the SSME and the RS-68, occurring in the past few decades.1 In recent years, however, there has been an increased interest in pursuing more effective launch propulsion technologies in the U.S., exemplified by the NASA Office of the Chief Technologist s inclusion of Launch Propulsion Systems as the first technological area in the Space Technology Roadmaps2. One area of particular interest to both government agencies and commercial entities has been the development of hydrocarbon engines; NASA and the Air Force Research Lab3 have expressed interest in the use of hydrocarbon fuels for their respective SLS Booster and Reusable Booster System concepts, and two major commercially-developed launch vehicles SpaceX s Falcon 9 and Orbital Sciences Antares feature engines that use RP-1 kerosene fuel. Compared to engines powered by liquid hydrogen, hydrocarbon-fueled engines have a greater propellant density (usually resulting in a lighter overall engine), produce greater propulsive force, possess easier fuel handling and loading, and for reusable vehicle concepts can provide a shorter turnaround time between launches. These benefits suggest that a hydrocarbon-fueled launch vehicle would allow for a cheap and frequent means of access to space.1 However, the time and money required for the development of a new engine still presents a major challenge. Long and costly design, development, testing and evaluation (DDT&E) programs underscore the importance of identifying critical technologies and prioritizing investment efforts. Trade studies must be performed on engine concepts examining the affordability, operability, and reliability of each concept, and quantifying the impacts of proposed technologies. These studies can be performed through use of the Technology Impact Forecasting (TIF) method. The Technology Impact Forecasting method is a normative forecasting technique that allows the designer to quantify the effects of adding new technologies on a given design. This method can be used to assess and identify the necessary technological improvements needed to close the gap that exists between the current design and one that satisfies all constraints imposed on the design. The TIF methodology allows for more design knowledge to be brought to the earlier phases of the design process, making use of tools such as Quality Function Deployments, Morphological Matrices, Response Surface Methodology, and Monte Carlo Simulations.2 This increased knowledge allows for more informed decisions to be made earlier in the design process, resulting in shortened design cycle time. This paper will investigate applying the TIF method, which has been widely used in aircraft applications, to the conceptual design of a hydrocarbon rocket engine. In order to reinstate a manned presence in space, the U.S. must develop an affordable and sustainable launch capability. Hydrocarbon-fueled rockets have drawn interest from numerous major government and commercial entities because they offer a low-cost heavy-lift option that would allow for frequent launches1. However, the development of effective new hydrocarbon rockets would likely require new technologies in order to overcome certain design constraints. The use of advanced design methods, such as the TIF method, enables the designer to identify key areas in need of improvement, allowing one to dial in a proposed technology and assess its impact on the system. Through analyses such as this one, a conceptual design for a hydrocarbon-fueled vehicle that meets all imposed requirements can be achieved.

Stuber, Eric↗
NASA NTRS2012

2011-2012 Dryden Center Innovation Fund End of the Year Report: Altitude-Compensating Rocket Nozzles

This report highlights one of the many successful projects at the NASA Dryden Flight Research Center that was approved for FY12 funding under the Center Innovation Fund. This project was focused on advancing the technology readiness level of one specific type of altitude-compensating nozzle: the dual-bell rocket nozzle. When considering a rocket's performance over its entire integrated trajectory, the dual-bell nozzle has been predicted to achieve a higher total impulse over the conventional bell nozzle, which is expected to result in a greater capability of payload mass to low-Earth orbit. Although the dual-bell rocket nozzle has been thoroughly studied for several decades, this nozzle has still not been adequately tested in a relevant flight-like environment. This report provides highlights and top-level details on the FY12 feasibility effort to advance this promising technology through flight test, a collaborative effort which leverages NASA Marshall's dual-bell nozzle research and development with Dryden's expertise in propulsion-focused flight testing. To accomplish this goal, the NASA F-15B is proposed as the testbed for the initial flight-test campaign to advance this greatly needed capability.

Jones, Daniel S.↗
NASA NTRS2012

National Institute for Rocket Propulsion Systems 1st Annual Workshop

The National Institute for Rocket Propulsion Systems (NIRPS) is a Government -wide initiative that seeks to ensure the resiliency of the Nation fs rocket propulsion community in order for the enterprise to remain vibrant and capable of providing reliable and affordable propulsion systems for the nation fs defense, civil and commercial needs. Recognizing that rocket propulsion is a multi-use technology that ensures the nation fs leadership in aerospace, the Government has a vested interest in maintaining this strategic capability through coordinated and synchronized acquisition programs and continual investments in research and development. NIRPS is a resource for collaboration and integration between all sectors of the U.S. propulsion enterprise, supporting policy development options, identifying technology requirements, and offering solutions that maximize national resources while ensuring that capability exists to meet future demand. NIRPS functions as a multi ]agency organization that our nation fs decision makers can look to for comprehensive information regarding all issues concerning the propulsion enterprise.

Doreswamy, Rajiv↗
NASA NTRS2004

Rocket Exhaust Cratering: Lessons Learned from Viking and Apollo

During the Apollo and Viking programs NASA expended considerable effort to study the cratering of the regolith when a rocket launches or lands on it. That research ensured the success of those programs but also demonstrated that cratering will be a serious challenge for other mission scenarios. Unfortunately, because three decades have elapsed since NASA last performed a successful retro-rocket landing on a large planetary body - and ironically because Apollo and Viking were successful at minimizing the cratering effects - the space agency has a minimized sense of the seriousness of the issue. The most violent phase of a cratering event is when the static overpressure of the rocket exhaust exceeds the bearing capacity of the soil. This bearing capacity failure (BCF) punches a small and highly concave cup into the surface. The shape of the cup then redirects the supersonic jet - along with a large flux of high-velocity debris - directly toward the spacecraft. This has been observed in terrestrial experiments but never quantified analytically. The blast from such an event will be more than just quantitatively greater than the cratering that occurred in the Apollo and Viking programs. It will be qualitatively different, because BCF had been successfully avoided in all those missions. In fact, the Viking program undertook a significant research and development effort and redesigned the spacecraft specifically for the purpose of avoiding BCF [1]. (See Figure 1.) Because the Apollo and Viking spacecraft were successful at avoiding those cratering effects, it was unnecessary to understand them. As a result, the physics of a BCF-driven cratering event have never been well understood. This is a critical gap in our knowledge because BCF is unavoidable in the Martian environment with the large landers necessary for human exploration, and in Lunar landings it must also be addressed because it may occur depending upon the design specifics of the spacecraft and the weakening of the regolity by gas diffusion.

Metzger, Philip T.↗
NASA NTRS2013

Proposed Flight Research of a Dual-Bell Rocket Nozzle Using the NASA F-15 Airplane

For more than a half-century, several types of altitude-compensating rocket nozzles have been proposed and analyzed, but very few have been adequately tested in a relevant flight environment. One type of altitude-compensating nozzle is the dual-bell rocket nozzle, which was first introduced into literature in 1949. Despite the performance advantages that have been predicted, both analytically and through static test data, the dual-bell nozzle has still not been adequately tested in a relevant flight environment. This paper proposes a method for conducting testing and research with a dual-bell rocket nozzle in a flight environment. We propose to leverage the existing NASA F-15 airplane and Propulsion Flight Test Fixture as the flight testbed, with the dual-bell nozzle operating during captive-carried flights, and with the nozzle subjected to a local flow field similar to that of a launch vehicle. The primary objective of this effort is not only to advance the technology readiness level of the dual-bell nozzle, but also to gain a greater understanding of the nozzle mode transitional sensitivity to local flow-field effects, and to quantify the performance benefits with this technology. The predicted performance benefits are significant, and may result in reducing the cost of delivering payloads to low-Earth orbit.

Jones, Daniel S.↗
NASA NTRS2013

Ground and Space-Based Measurement of Rocket Engine Burns in the Ionosphere

On-orbit firings of both liquid and solid rocket motors provide localized disturbances to the plasma in the upper atmosphere. Large amounts of energy are deposited to ionosphere in the form of expanding exhaust vapors which change the composition and flow velocity. Charge exchange between the neutral exhaust molecules and the background ions (mainly O+) yields energetic ion beams. The rapidly moving pickup ions excite plasma instabilities and yield optical emissions after dissociative recombination with ambient electrons. Line-of-sight techniques for remote measurements rocket burn effects include direct observation of plume optical emissions with ground and satellite cameras, and plume scatter with UHF and higher frequency radars. Long range detection with HF radars is possible if the burns occur in the dense part of the ionosphere. The exhaust vapors initiate plasma turbulence in the ionosphere that can scatter HF radar waves launched from ground transmitters. Solid rocket motors provide particulates that become charged in the ionosphere and may excite dusty plasma instabilities. Hypersonic exhaust flow impacting the ionospheric plasma launches a low-frequency, electromagnetic pulse that is detectable using satellites with electric field booms. If the exhaust cloud itself passes over a satellite, in situ detectors measure increased ion-acoustic wave turbulence, enhanced neutral and plasma densities, elevated ion temperatures, and magnetic field perturbations. All of these techniques can be used for long range observations of plumes in the ionosphere. To demonstrate such long range measurements, several experiments were conducted by the Naval Research Laboratory including the Charged Aerosol Release Experiment, the Shuttle Ionospheric Modification with Pulsed Localized Exhaust experiments, and the Shuttle Exhaust Ionospheric Turbulence Experiments.

Bernhardt, P. A.↗
NASA NTRS2013

Proposed Flight Research of a Dual-Bell Rocket Nozzle Using the NASA F-15 Airplane

For more than a half-century, several types of altitude-compensating rocket nozzles have been proposed and analyzed, but very few have been adequately tested in a relevant flight environment. One type of altitude-compensating nozzle is the dual-bell rocket nozzle, which was first introduced into literature in 1949. Despite the performance advantages that have been predicted, both analytically and through static test data, the dual-bell nozzle has still not been adequately tested in a relevant flight environment. This paper proposes a method for conducting testing and research with a dual-bell rocket nozzle in a flight environment. We propose to leverage the existing NASA F-15 airplane and Propulsion Flight Test Fixture as the flight testbed, with the dual-bell nozzle operating during captive-carried flights, and with the nozzle subjected to a local flow field similar to that of a launch vehicle. The primary objective of this effort is not only to advance the technology readiness level of the dual-bell nozzle, but also to gain a greater understanding of the nozzle mode transitional sensitivity to local flow-field effects, and to quantify the performance benefits with this technology. The predicted performance benefits are significant, and may result in reducing the cost of delivering payloads to low-Earth orbit.

Jones, Daniel S.↗
NASA NTRS2014

Coupled Fluid-Structure Interaction Analysis of Solid Rocket Motor with Flexible Inhibitors

Flexible inhibitors are generally used in solid rocket motors (SRMs) as a means to control the burning of propellant. Vortices generated by the flow of propellant around the flexible inhibitors have been identified as a driving source of instabilities that can lead to thrust oscillations in launch vehicles. Potential coupling between the SRM thrust oscillations and structural vibration modes is an important risk factor in launch vehicle design. As a means to predict and better understand these phenomena, a multidisciplinary simulation capability that couples the NASA production CFD code, Loci/CHEM, with CFDRC's structural finite element code, CoBi, has been developed. This capability is crucial to the development of NASA's new space launch system (SLS). This paper summarizes the efforts in applying the coupled software to demonstrate and investigate fluid-structure interaction (FSI) phenomena between pressure waves and flexible inhibitors inside reusable solid rocket motors (RSRMs). The features of the fluid and structural solvers are described in detail, and the coupling methodology and interfacial continuity requirements are then presented in a general Eulerian-Lagrangian framework. The simulations presented herein utilize production level CFD with hybrid RANS/LES turbulence modeling and grid resolution in excess of 80 million cells. The fluid domain in the SRM is discretized using a general mixed polyhedral unstructured mesh, while full 3D shell elements are utilized in the structural domain for the flexible inhibitors. Verifications against analytical solutions for a structural model under a steady uniform pressure condition and under dynamic modal analysis show excellent agreement in terms of displacement distribution and eigenmode frequencies. The preliminary coupled results indicate that due to acoustic coupling, the dynamics of one of the more flexible inhibitors shift from its first modal frequency to the first acoustic frequency of the solid rocket motor. This insight could have profound implications for SRM and flexible inhibitor designs for current and future launch vehicles including SLS.

Yang, H. Q.↗
NASA NTRS2014

Designing Liquid Rocket Engine Injectors for Performance, Stability, and Cost

NASA is developing the Space Launch System (SLS) for crewed exploration missions beyond low Earth orbit. Marshall Space Flight Center (MSFC) is designing rocket engines for the SLS Advanced Booster (AB) concepts being developed to replace the Shuttle-derived solid rocket boosters. One AB concept uses large, Rocket-Propellant (RP)-fueled engines that pose significant design challenges. The injectors for these engines require high performance and stable operation while still meeting aggressive cost reduction goals for access to space. Historically, combustion stability problems have been a critical issue for such injector designs. Traditional, empirical injector design tools and methodologies, however, lack the ability to reliably predict complex injector dynamics that often lead to combustion stability. Reliance on these tools alone would likely result in an unaffordable test-fail-fix cycle for injector development. Recently at MSFC, a massively parallel computational fluid dynamics (CFD) program was successfully applied in the SLS AB injector design process. High-fidelity reacting flow simulations were conducted for both single-element and seven-element representations of the full-scale injector. Data from the CFD simulations was then used to significantly augment and improve the empirical design tools, resulting in a high-performance, stable injector design.

Westra, Douglas G.↗
NASA NTRS2014

One Dimensional Analysis Model of a Condensing Spray Chamber Including Rocket Exhaust Using SINDA/FLUINT and CEA

Modeling droplet condensation via CFD codes can be very tedious, time consuming, and inaccurate. CFD codes may be tedious and time consuming in terms of using Lagrangian particle tracking approaches or particle sizing bins. Also since many codes ignore conduction through the droplet and or the degradating effect of heat and mass transfer if noncondensible species are present, the solutions may be inaccurate. The modeling of a condensing spray chamber where the significant size of the water droplets and the time and distance these droplets take to fall, can make the effect of droplet conduction a physical factor that needs to be considered in the model. Furthermore the presence of even a relatively small amount of noncondensible has been shown to reduce the amount of condensation. It is desirable then to create a modeling tool that addresses these issues. The path taken to create such a tool is illustrated. The application of this tool and subsequent results are based on the spray chamber in the Spacecraft Propulsion Research Facility (B2) located at NASA's Plum Brook Station that tested an RL-10 engine. The platform upon which the condensation physics is modeled is SINDAFLUINT. The use of SINDAFLUINT enables the ability to model various aspects of the entire testing facility, including the rocket exhaust duct flow and heat transfer to the exhaust duct wall. The ejector pumping system of the spray chamber is also easily implemented via SINDAFLUINT. The goal is to create a transient one dimensional flow and heat transfer model beginning at the rocket, continuing through the condensing spray chamber, and finally ending with the ejector pumping system. However the model of the condensing spray chamber may be run independently of the rocket and ejector systems detail, with only appropriate mass flow boundary conditions placed at the entrance and exit of the condensing spray chamber model. The model of the condensing spray chamber takes into account droplet conduction as well as the degrading effect of mass and heat transfer due to the presence of noncondensibles. The one dimension model of the condensing spray chamber makes no presupposition on the pressure profile within the chamber, allowing the implemented droplet physics of heat and mass transfer coupled to the SINDAFLUINT solver to determine a transient pressure profile of the condensing spray chamber. Model results compare well to the RL-10 engine pressure test data.

Chemical Equilibrium↗
NASA NTRS2014

One Dimensional Analysis Model of a Condensing Spray Chamber Including Rocket Exhaust Using SINDA/FLUINT and CEA

Modeling droplet condensation via CFD codes can be very tedious, time consuming, and inaccurate. CFD codes may be tedious and time consuming in terms of using Lagrangian particle tracking approaches or particle sizing bins. Also since many codes ignore conduction through the droplet and or the degradating effect of heat and mass transfer if noncondensible species are present, the solutions may be inaccurate. The modeling of a condensing spray chamber where the significant size of the water droplets and the time and distance these droplets take to fall, can make the effect of droplet conduction a physical factor that needs to be considered in the model. Furthermore the presence of even a relatively small amount of noncondensible has been shown to reduce the amount of condensation [Ref 1]. It is desirable then to create a modeling tool that addresses these issues. The path taken to create such a tool is illustrated. The application of this tool and subsequent results are based on the spray chamber in the Spacecraft Propulsion Research Facility (B2) located at NASA's Plum Brook Station that tested an RL-10 engine. The platform upon which the condensation physics is modeled is SINDAFLUINT. The use of SINDAFLUINT enables the ability to model various aspects of the entire testing facility, including the rocket exhaust duct flow and heat transfer to the exhaust duct wall. The ejector pumping system of the spray chamber is also easily implemented via SINDAFLUINT. The goal is to create a transient one dimensional flow and heat transfer model beginning at the rocket, continuing through the condensing spray chamber, and finally ending with the ejector pumping system. However the model of the condensing spray chamber may be run independently of the rocket and ejector systems detail, with only appropriate mass flow boundary conditions placed at the entrance and exit of the condensing spray chamber model. The model of the condensing spray chamber takes into account droplet conduction as well as the degrading effect of mass and heat transfer due to the presence of noncondensibles. The one dimension model of the condensing spray chamber makes no presupposition on the pressure profile within the chamber, allowing the implemented droplet physics of heat and mass transfer coupled to the SINDAFLUINT solver to determine a transient pressure profile of the condensing spray chamber. Model results compare well to the RL-10 engine pressure test data.

Condensing Spray Model↗
NASA NTRS2013

Rocket Sled Propelled Testing of a Supersonic Inflatable Aerodynamic Decelerator

Decelerators (IADs) have traditionally been tested in wind tunnels. As the limitations of these test facilities are reached, other avenues must be pursued. The IAD being tested is a Supersonic IAD (SIAD), which attaches just aft of the heatshield around the perimeter of an entry body. This 'attached torus' SIAD is meant to improve the accuracy of landing for robotic class missions to Mars and allow for potentially increased payloads. The SIAD Design Verification (SDV) test aims to qualify the SIAD by applying a targeted aerodynamic load to the vehicle. While many test architectures were researched, a rocket sled track was ultimately chosen to be the most cost effective way to achieve the desired dynamic pressures. The Supersonic Naval Ordnance Research Track (SNORT) at the Naval Air Warfare Center Weapons Division (NAWCWD) China Lake is a four mile test track, traditionally used for warhead and ejection seat testing. Prior to SDV, inflatable drag bodies have been tested on this particular track. Teams at Jet Propulsion Laboratory (JPL) and NAWCWD collaborate together to design and fabricate one of the largest sleds ever built. The SDV sled is comprised of three individual sleds: a Pusher Sled which holds the solid booster rockets, an Item Sled which supports the test vehicle, and a Camera Sled that is pushed in front for in-situ footage and measurements. The JPL-designed Test Vehicle has a full-scale heatshield shape and contains all instrumentation and inflation systems necessary to inflate and test a SIAD. The first campaign that is run at SNORT tested all hardware and instrumentation before the SIAD was ready to be tested. For each of the three tests in this campaign, the number of rockets and top speed was increased and the data analyzed to ensure the hardware is safe at the necessary accelerations and aerodynamic loads.

Supersonic Inflatable Aerodynamic Decelerators (SI↗