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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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161 records · Page 9
All sourcesNASA NTRSDOE OSTI
NASA NTRS2013

Entry, Descent, and Landing with Propulsive Deceleration: Supersonic Retropropulsion Wind Tunnel Testing and Shock Phenomena

The future exploration of the Solar System will require innovations in transportation and the use of entry, descent, and landing (EDL) systems at many planetary landing sites. The cost of space missions has always been prohibitive, and using the natural planetary and planet's moon atmospheres for entry, and descent can reduce the cost, mass, and complexity of these missions. This paper will describe some of the EDL ideas for planetary entry and survey the overall technologies for EDL that may be attractive for future Solar System missions. Future EDL systems may include an inflatable decelerator for the initial atmospheric entry and an additional supersonic retro-propulsion (SRP) rocket system for the final soft landing. A three engine retro-propulsion configuration with a 2.5 inch diameter sphere-cone aeroshell model was tested in the NASA Glenn 1x1 Supersonic Wind Tunnel (SWT). The testing was conducted to identify potential blockage issues in the tunnel, and visualize the rocket flow and shock interactions during supersonic and hypersonic entry conditions. Earlier experimental testing of a 70 degree Viking-like (sphere-cone) aeroshell was conducted as a baseline for testing of a supersonic retro-propulsion system. This baseline testing defined the flow field around the aeroshell and from this comparative baseline data, retro-propulsion options will be assessed. Images and analyses from the SWT testing with 300- and 500-psia rocket engine chamber pressures are presented here. In addition, special topics of electromagnetic interference with retro-propulsion induced shock waves and retro-propulsion for Earth launched booster recovery are also addressed.

Palaszewski, Bryan↗
NASA NTRS2015

Orbit Determination for the Lunar Reconnaissance Orbiter Using an Extended Kalman Filter

Since launch, the FDF has performed daily OD for LRO using the Goddard Trajectory Determination System (GTDS). GTDS is a batch least-squares (BLS) estimator. The tracking data arc for OD is 36 hours. Current operational OD uses 200 x 200 lunar gravity, solid lunar tides, solar radiation pressure (SRP) using a spherical spacecraft area model, and point mass gravity for the Earth, Sun, and Jupiter. LRO tracking data consists of range and range-rate measurements from: Universal Space Network (USN) stations in Sweden, Germany, Australia, and Hawaii. A NASA antenna at White Sands, New Mexico (WS1S). NASA Deep Space Network (DSN) stations. DSN data was sparse and not included in this study. Tracking is predominantly (50) from WS1S. The OD accuracy requirements are: Definitive ephemeris accuracy of 500 meters total position root-mean-squared (RMS) and18 meters radial RMS. Predicted orbit accuracy less than 800 meters root sum squared (RSS) over an 84-hour prediction span.

LRO↗
NASA NTRS2016

A High-Heritage Blunt-Body Entry, Descent, and Landing Concept for Human Mars Exploration

Human-scale landers require the delivery of much heavier payloads to the surface of Mars than is possible with entry, descent, and landing (EDL) approaches used to date. A conceptual design was developed for a 10 m diameter crewed Mars lander with an entry mass of approx.75 t that could deliver approx.28 t of useful landed mass (ULM) to a zero Mars areoid, or lower, elevation. The EDL design centers upon use of a high ballistic coefficient blunt-body entry vehicle and throttled supersonic retro-propulsion (SRP). The design concept includes a 26 t Mars Ascent Vehicle (MAV) that could support a crew of 2 for approx.24 days, a crew of 3 for approx.16 days, or a crew of 4 for approx.12 days. The MAV concept is for a fully-fueled single-stage vehicle that utilizes a single pump-fed 250 kN engine using Mono-Methyl Hydrazine (MMH) and Mixed Oxides of Nitrogen (MON-25) propellants that would deliver the crew to a low Mars orbit (LMO) at the end of the surface mission. The MAV concept could potentially provide abort-to-orbit capability during much of the EDL profile in response to fault conditions and could accommodate return to orbit for cases where the MAV had no access to other Mars surface infrastructure. The design concept for the descent stage utilizes six 250 kN MMH/MON-25 engines that would have very high commonality with the MAV engine. Analysis indicates that the MAV would require approx.20 t of propellant (including residuals) and the descent stage would require approx.21 t of propellant. The addition of a 12 m diameter supersonic inflatable aerodynamic decelerator (SIAD), based on a proven flight design, was studied as an optional method to improve the ULM fraction, reducing the required descent propellant by approx.4 t.

Mars lander↗
NASA NTRS2016

Engineering Considerations Applied to Starshade Repointing

Engineering analysis has been carried out on orbit dynamics that drive the delta-v budget for repointing a free-flying starshade occulter for viewing exoplanets with a space telescope. This analysis has application to the design of starshade spacecraft and yield calculations of observations of exoplanets using a space telescope and a starshade. Analysis was carried out to determine if there may be some advantage for the global delta-v budget if the telescope performs orbit changing delta-v maneuvers as part of the telescope-starshade alignment for observing exoplanets. Analysis of the orbit environmental forces at play found no significant advantage in having the telescope participate in delta-v maneuvers for exoplanet observation repointing. A separate analysis of starshade delta-v for repointing found that the orbit dynamics of the starshade is driven by multiple simultaneous variables that need to be considered together in order to create an effective estimate of delta-v over an exoplanet observation campaign. These include area of the starshade, dry mass of the starshade spacecraft, and propellant mass of the starshade spacecraft. Solar radiation pressure has the potential to play a dominant role in the orbit dynamics and delta-v budget. SRP effects are driven by the differences in the mass, area, and coefficients of reflectivity of the observing telescope and the starshade. The propellant budget cannot be effectively estimated without a conceptual design of a starshade spacecraft including the propulsion system. The varying propellant mass over the mission is a complexity that makes calculating the propellant budget less straightforward.

delta-V↗
NASA NTRS2017

Application of IRTAM to Support ISS Program Safety

The International Space Station (ISS) orbits near the F-peak of the ionosphere (approximately 400 km altitude). Generally, satellites orbiting at this altitude would have a floating potential (FP) of approximately -1 V due to the electron temperature (Te). However, the ISS has 8 large negatively grounded 160 V solar array wings (SAW) that collect a significant electron current from the ionosphere. This current drives the ISS FP much more negative during insolation and is highly dependent on the electron density (Ne). Also, due to the size of the ISS, magnetic inductance caused by the geomagnetic field produces a delta potential up to 40 V across the truss, possibly producing positive potentials. During Extravehicular Activity (EVA) the negative FP can lead to an arcing hazard when it exceeds -45.5 V, and the positive FP can produce a DC current high enough to stimulate the astronaut's muscles and also cause a hazard. Data collected from the Floating Potential Monitoring Unit (FPMU) have shown that the probability of either of these hazards occurring during times with quiet to moderately disturbed geomagnetic activity is low enough to no longer be considered a risk. However, a study of the ionosphere Ne during severe geomagnetic storm activity has shown that the Ne can be enhanced by a factor of 6 in the ISS orbit. As a result, the ISS Safety Review Panel (SRP) requires that ionospheric conditions be monitored using the FPMU in conjunction with the ISS Plasma Interaction Model (PIM) to determine if a severe geomagnetic storm could result in a plasma environment that could produce a hazard. A 'Real-Time' plasma hazard assessment process was developed to support ISS Program real-time decision making providing constraint relief information for EVAs planning and operations. This process incorporates 'real time' ionospheric conditions, ISS solar arrays' orientation, ISS flight attitude, and where the EVA will be performed on the ISS. This assessment requires real time data that is presently provided by the FPMU including ISS floating potential, along with ionospheric Ne and Te, in order to determine the present environment. Once the present environment conditions are known to be either above, below, or near the current IRI values, the IRI is used to forecast what the environment could become in the event of a severe geomagnetic storm. If the FPMU should fail, the Space Environments team needs another source of data which is utilized to support a short-term forecast for EVAs. The IRI Real-Time Assimilative Mapping (IRTAM) model is an ionospheric model that uses real time measurements from approximately 70 digisondes to produce foF2 and hmF2 global maps in 15 minute cadence. The Boeing Space Environments team has used the IRI coefficients produced in IRTAM to calculate the Ne along the ISS orbital track. The results of the IRTAM model have been compared to FPMU measurements and show excellent agreement (figure 1). IRTAM has been identified as a potential FPMU back-up system will be used as a backup for the FPMU to support the ISS Program following completion of an FPMU/IRTAM validation campaign.

Hartman, William A.↗
NASA NTRS2016

A High-Heritage Blunt-Body Entry, Descent, and Landing Concept for Human Mars Exploration

Human-scale landers require the delivery of much heavier payloads to the surface of Mars than is possible with entry, descent, and landing (EDL) approaches used to date. A conceptual design was developed for a 10 m diameter crewed Mars lander with an entry mass of approx. 75 t that could deliver approx. 28 t of useful landed mass (ULM) to a zero Mars areoid, or lower, elevation. The EDL design centers upon use of a high ballistic coefficient blunt-body entry vehicle and throttled supersonic retro-propulsion (SRP). The design concept includes a 26 t Mars Ascent Vehicle (MAV) that could support a crew of 2 for approx. 24 days, a crew of 3 for approx.16 days, or a crew of 4 for approx.12 days. The MAV concept is for a fully-fueled single-stage vehicle that utilizes a single pump-fed 250 kN engine using Mono-Methyl Hydrazine (MMH) and Mixed Oxides of Nitrogen (MON-25) propellants that would deliver the crew to a low Mars orbit (LMO) at the end of the surface mission. The MAV concept could potentially provide abort-to-orbit capability during much of the EDL profile in response to fault conditions and could accommodate return to orbit for cases where the MAV had no access to other Mars surface infrastructure. The design concept for the descent stage utilizes six 250 kN MMH/MON-25 engines that would have very high commonality with the MAV engine. Analysis indicates that the MAV would require approx. 20 t of propellant (including residuals) and the descent stage would require approx. 21 t of propellant. The addition of a 12 m diameter supersonic inflatable aerodynamic decelerator (SIAD), based on a proven flight design, was studied as an optional method to improve the ULM fraction, reducing the required descent propellant by approx.4 t.

Mars lander↗
NASA NTRS2018

Entry, Descent, and Landing Performance for a Mid-Lift-to-Drag Ratio Vehicle at Mars

In an effort to mature the design of the Mid-Lift-to-Drag ratio Rigid Vehicle (MRV) candidate of the NASA Evolvable Mars Campaign (EMC) architecture study, end-to-end six-degree-of-freedom (6DOF) simulations are needed to ensure a successful entry, descent, and landing (EDL) design. The EMC study is assessing different vehicle and mission architectures to determine which candidate would be best to deliver a 20 metric ton payload to the surface of Mars. Due to the large mass payload and the relatively low atmospheric density of Mars, all candidates of the EMC study propose to use Supersonic Retro-Propulsion (SRP) throughout the descent and landing phase, as opposed to parachutes, in order to decelerate to a subsonic touchdown. This paper presents a 6DOF entry-to-landing performance and controllability study with sensitivities to dispersions, particularly in the powered descent and landing phases.

mars↗
NASA NTRS2018

Dismantling Rubble Pile Asteroids with AoES (Area-of-Effect Soft-bots)

Area-of-Effect Softbots (AoES) are soft-robotic spacecraft that are designed with a large, flexible surface area to leverage the dynamical environment at rubble pile asteroids. In particular, this surface area allows AoES to use adhesive forces, both naturally arising from van der Waals forces between the AoES and the asteroid regolith, and by using active electroadhesion, as well as using SRP forces to provide fuel free orbit and hopping trajectory control. The main purpose of the bus structure is to house a digging and launching mechanism that can liberate and launch asteroid regolith off the surface of the asteroid to be collected in orbit.

McMahon, Jay↗
NASA NTRS2022

Pre-Installation Acceptance (PIA) Functional Performance of the Design Verification Test (DVT) Exploration Extra-vehicular Mobility Unit (xEMU)

In an effort that began with technology investment by NASA in a few key components during the Constellation Program and then evolved to demonstrate a packaged Portable Life Support System (PLSS) as part of the Advanced Exploration Systems (AES) Program, the next evolution of the PLSS is now a key component of the Exploration Extra-vehicular Mobility Unit (xEMU) and is assembled as a Design Verification Test (DVT) unit. The xEMU has been detailed with respect to completing a demonstration on the International Space Station (ISS) with support of units for initial lunar capability. The xEMU completed the Preliminary Design Review (PDR) with subsequent Safety Review Panel (SRP) Phase I reviews in 2019-2020. The objectives for DVT are to validate requirements, train the team, learn how to fabricate the hardware with appropriate process controls, assemble the hardware, test the hardware, determine the failure mechanisms/limits of the hardware design and buy down the most risk possible for the qualification and flight phases of the development. With completion of the assembly and initial functional testing of the DVT PLSS in laboratory ambient conditions and vacuum conditions the xEMU has progressed significantly into the DVT objectives. A key part of the test sequences for DVT and all future phases is the Pre-Installation Acceptance (PIA) functional testing which validates the performance of integrated systems including: primary thermal control, auxiliary thermal control, suit ventilation, primary oxygen, secondary oxygen, power distribution, as well as caution and warning all with respect to the applied requirements. This discussion will include an overview of the assembly, summary of the PIA functional testing, lessons learned, and corrective actions implemented moving forward into the remainder of the DVT phase for xEMU.

Colin Campbell↗
NASA NTRS2022

Permutations of the Exploration Extra-vehicular Mobility Unit (xEMU) for Mission Specific Objectives

The Advanced Portable Life Support System began first with NASA investment during the Constellation Program focused on new technologies that were central to the new architecture but with a consideration for development of the most flexible and adaptable design to address the varied future needs for Extra-Vehicular Activity (EVA) capability by the agency as part of the exploration of the moon and beyond. Since that time, the Portable Life Support System (PLSS) architecture has been demonstrated at bread-board level, packaged prototype level with Human In-The Loop (HITL) testing, and has now matured into what has become known as the Exploration Extravehicular Mobility Unit (xEMU) PLSS focused on a demonstration on the International Space Station (ISS) and support for initial lunar missions. The xEMU completed Preliminary Design Review (PDR) and the associated Safety Review Panel (SRP) Phase I reviews in 2019-2020 and is progressing through the Design Verification Test (DVT) phase in 2021. As the agency updates planning for lunar both near term, long term, and beyond, there are many possible permutations of the existing architecture that can be accommodated to enable varied mission objectives. This paper will discuss the current baseline DVT design for xEMU as well as potential permutations of the PLSS architecture to accomplish an Apollo style short-duration down and out mission, a long term lunar ultra-lightweight approach, as well as a potential SuitPort accommodating architecture.

Colin Campbell↗
NASA NTRSDate not supplied

Preparations for Tomographic Background-Oriented Schlieren at the 31-Inch Mach 10 Wind Tunnel

This paper details effort by researchers at NASA Langley Research Center to prepare for an upcoming high-speed tomographic background-oriented schlieren test entry in the 31-Inch Mach 10 wind tunnel. In order to troubleshoot potential problems that will arise when setting up the experiment in the actual test facility, a full-scale mock-up of the 31-Inch Mach 10 wind tunnel test section was constructed in the laboratory to provide equivalent characteristic optical access and mounting options. A 3D-printed model that replicates the scale of the forebody of a supersonic retropropulsion (SRP) model that will be used in the final test entry has been fabricated with a central nozzle that can be connected to a high-pressure gas supply to approximate the nozzle flow that will occur. Six high-speed imaging systems consisting of a high-speed camera with a fiber illuminator module have been assembled and mounted on the test section mock-up to provide six unique perspective views of the nozzle flow from the model. Tests with one of the imaging systems using a pulsed laser light source showed that sufficient signal intensity and illumination uniformity was achieved, as demonstrated by the 2D BOS data obtained with the system. Additional work to show that retroreflective material would survive inside of the test section of the 31-Inch Mach 10 tunnel was also performed. Finally, a remotely-controlled tomographic BOS calibration system is discussed.

Brett F Bathel↗
NASA NTRS2024

Comparison of OVERFLOW Computational and Experimental Results of the CobraMRV Mars Entry Vehicle Concept during Supersonic Retropropulsion

The CFD solver OVERFLOW was used to simulate the CobraMRV undergoing supersonic retropropulsion (SRP) in the Langley Unitary Plan Wind Tunnel as part of a CFD and wind-tunnel integration effort. The aerodynamics of the CobraMRV are summarized and comparisons to experimental data at flight conditions of interest are presented. The sensitivity of vehicle loads and flow field unsteadiness to CFD parameters at a subset of operating conditions are investigated.

CFD↗
NASA NTRSDate not supplied

Collaborative Host Facility for Lunar Operations

This project establishes a new collaborative simulation connectivity capability for JSC utilizing existing facilities. The principal collaborating organizations are Safety & Mission Assurance (Code NA) and Engineering Directorate (Code EA), but the capability is open for any JSC user. The core of the connectivity resides in Code ER7’s Concept Exploration Laboratory (CEL) and is part of ER7’s Systems Engineering Simulator (SES) complex. NA Subject Matter Experts (SMEs) work proactively in situ with EA and commercial crew counterparts to evaluate Safety Review Panel (SRP) protocols for landers, habitation modules, and lunar roving systems using the same tools, terrain models and data as the designers. NA software initiated under two JSC Center Innovation Fund (CIF) and Innovation Charge account (ICA) funded projects will be used to configure the CEL to function as an EA maintained facility. Collaboration via the CEL with the U.S. Space Force (USSF) provides visibility for USSF. The potential for commercial crew and vehicle suppliers will require IT security and associated security architecture to be implemented in additional CIF based proposals.

Collaborative Simulation Hosting↗
NASA NTRSDate not supplied

Comparison of OVERFLOW Computational and Experimental Results of the CobraMRV Mars Entry Vehicle Concept during Supersonic Retropropulsion

The CFD solver Overflow was used to simulate the CobraMRV undergoing supersonic retropropulsion (SRP) in the Langley Unitary Plan Wind Tunnel as part of a CFD and wind tunnel integration effort. Selection of numerical methods and resulting comparisons to selected tunnel conditions of interest are detailed. Comparisons of mean pressures with static taps and pressure-sensitive paint are provided for multiple SST-based turbulence models. URANS predicitons of heat shield pressures at high thrust conditions are poor compared to TMRC and DES predictions. Comparisons to experimental measurements of pressure fluctuations indicate complex relationships between heat shield unsteadiness, flight Mach number, and engine thrust coefficient. Bow shock topology and shock cell structure between experimental and OVERFLOW data show good agreement in Schlieren imaging. Recommendations for future investigation into flow field unsteadiness and turbulence models are suggested.

CFD↗
NASA NTRSDate not supplied

Computational Investigation of Retropropulsion Operating Environments with a GPU-Enabled Detached Eddy Simulation Approach

Human exploration of the surface of Mars will require an extended powered descent phase of flight, during which aerodynamic-propulsive interference effects can be significant. Characterization of these environments to enable implementation of this technology into a flight vehicle will rely heavily on computational simulation. This work advances the understanding of retropropulsion aerodynamics through application of a massively parallel detached eddy simulation approach on a GPU-accelerated computational framework, yielding data that are largely unachievable with conventional high-performance computing resources. This work includes time-dependent and time-averaged forces and moments on a conceptual, full-scale vehicle in environments and operating conditions relevant to human Mars exploration. Conditions are examined where the engine exhaust flow transitions between over-expanded and under-expanded flow structures, and flight operation will require the ability to maintain control of the vehicle during such a transition. These transitions occur as the vehicle decelerates, and as such, this investigation includes supersonic, transonic, and subsonic flight conditions. Options for vehicle control during powered flight include differential throttling of the engines. This paper provides an overview of the computational campaign, approach, and discussion of results in characterizing the resulting aerodynamics for differential throttling with retropropulsion in atmospheric environments.

EDL↗