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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 NTRS2018

TPSAS-NF1676L-31282-DND

Several technology investments, beyond those used for robotic missions, are required to develop Mars human scale Entry, Descent, and Landing (EDL) systems. In a resource-constrained environment, studies play the critical role of identifying the most feasible technical paths and high payoff investments. One such NASA multi-directorate, multi-center study, is called the Entry, Descent and Landing Architecture Study (EDLAS). This paper presents a summary of phase two of the study conducted from October 2016 to September 2017. Study ground rules and assumptions are provided by NASA?s Human Exploration and Operations Mission Directorate (HEOMD) and include a Mars architecture and Mars surface lander payload manifests. Four unique entry technologies with the potential to deliver a specified 20 t human scale payload to the Mars surface are considered for analysis. Two of these techniques are evolutionary rigid vehicles, derived from robotic capsule missions and the space shuttle. Likewise, two revolutionary deployable vehicles are considered, the Adaptable Deployable Entry Placement Technology (ADEPT) and Hypersonic Inflatable Aerodynamic Decelerator (HIAD). This paper summarizes updates to the entry technology designs, the analysis motivation and approach, figures of merit by which the configurations are compared, and results of the analyses. Finally, findings are presented with the recommendation that two of the four configurations, one evolutionary and one revolutionary technology, continue to be studied in future human scale EDL studies.

Alicia Dwyer Cianciolo↗
NASA NTRS2016

Human Mars Entry, Descent and Landing Architectures Study Overview

Landing humans on Mars will require entry, descent and landing (EDL) capability beyond the current state of the art. Nearly twenty times more delivered payload and an order of magnitude improvement in precision landing capability will be necessary. Several EDL technologies capable of meeting the human class payload delivery requirements are being considered. The EDL technologies considered include low lift-to-drag vehicles like Hypersonic Inflatable Aerodynamic Decelerators (HIAD), Adaptable Deployable Entry and Placement Technology (ADEPT), and mid range lift-to-drag vehicles like rigid aeroshell configurations. To better assess EDL technology options and sensitivities to future human mission design variations, a series of design studies has been conducted. The design studies incorporate EDL technologies with conceptual payload arrangements defined by the Evolvable Mars Campaign to evaluate the integrated system with higher fidelity than have been performed to date. This paper describes the results of the design studies for a lander design using the HIAD, ADEPT and rigid shell entry technologies and includes system and subsystem design details including mass and power estimates. This paper will review the point design for three entry configurations capable of delivering a 20 t human class payload to the surface of Mars.

Polsgrove, Tara T.↗
NASA NTRS2015

Adaptable, Deployable Entry and Placement Technology (ADEPT) Overview of FY15 Accomplishments

ADEPT is an atmospheric entry architecture for missions to most planetary bodies with atmospheres: Current Technology development project funded under STMD Game Changing Development Program (FY12 start); stowed inside the launch vehicle shroud and deployed in space prior to entry; low ballistic coefficient (less than 50 kilograms per square meter) provides a benign deceleration and thermal environment to the payload; High-temperature ribs support three dimensional woven carbon fabric to generate drag and withstand high heating.

technology↗
NASA NTRSDate not supplied

Design and Qualification or the Low-Earth Orbit Flight Test of an Inflatable Decelerator(LOFTID) Payload Adapter Separation System (PASS)

On November 10, 2022, NASA and United Launch Alliance (ULA) launched Low-Earth Orbit Flight Test of an Inflatable Decelerator (LOFTID) as a secondary payload on an Atlas V Centaur from Vandenburg Space Force Base (VSFB). After deploying the primary payload, JPSS-2, to a sun-synchronous orbit, the Centaur reoriented LOFTID for reentry, ejecting the Payload Adaptor to expose the stowed LOFTID reentry vehicle (RV). The Hypersonic Inflatable Aerodynamic Decelerator (HIAD) was then deployed and inflated, and the Centaur reoriented the RV to the desired entry attitude, spun the vehicle to roughly three rpm, and separated the reentry vehicle. After nearly half an orbit of free flight, LOFTID reentered the atmosphere at >8 km/sec, demonstrating stable flight from hypersonic entry through subsonic parachute deployment. A mission-unique Payload Adapter Separation System (PASS) enabled the successful LOFTID flight demonstration, allowing the launch vehicle to accommodate a superstack of two independent, similarly sized payloads. This paper discusses the design, development, and qualification of the LOFTID PASS along with lessons learned.

HIAD↗
NASA NTRSDate not supplied

LOFTID (Low-Earth Orbit Flight Test of an Inflatable Decelerator) PASS (Payload Adapter Separation System) Design & Qualification

On November 10, 2022, NASA, in partnership with United Launch Alliance (ULA), launched Low-earth Orbit Flight Test of an Inflatable Decelerator (LOFTID) as a secondary payload on an Atlas V Centaur out of the Vandenburg Space Force Base (VSFB). After successfully delivering the primary payload, Joint Polar Satellite System-2 (JPSS-2), to a sun synchronous trajectory, the Centaur upper stage reoriented LOFTID onto the desired reentry trajectory. After conducting a de-orbit burn to enter the atmosphere the Payload Adaptor was ejected to expose the packed LOFTID vehicle. The LOFTID Hypersonic Inflatable Aerodynamic Decelerator (HIAD) was deployed and inflated. The Centaur pointed LOFTID to the desired entry attitude and spun the vehicle up to roughly three rpm before separating the reentry vehicle over the Middle East. The LOFTID vehicle flew freely before reentering the atmosphere over Alaska at >8km/sec and decelerating as designed. LOFTID demonstrated stable flight from hypersonic entry through subsonic parachute deployment. LOFTID was enabled by a mission-unique Payload Adapter Separation System (PASS) which separated the Payload Adapter prior to the start of the LOFTID flight demonstration, allowing the launch vehicle to accommodate a superstack of two independent, similarly sized payloads. This paper will discuss the design, development, and qualification effort of the LOFTID PASS.

Sean M Hancock↗
NASA NTRS2010

Orion Entry Display Feeder and Interactions with the Entry Monitor System

The Orion spacecraft is designed to return astronauts to a landing within 10 km of the intended landing target from low Earth orbit, lunar direct-entry, and lunar skip-entry trajectories. Al pile the landing is nominally controlled autonomously, the crew can fly precision entries manually in the event of an anomaly. The onboard entry displays will be used by the crew to monitor and manually fly the entry, descent, and landing, while the Entry Monitor System (EMS) will be used to monitor the health and status of the onboard guidance and the trajectory. The entry displays are driven by the entry display feeder, part of the Entry Monitor System (EMS). The entry re-targeting module, also part of the EMS, provides all the data required to generate the capability footprint of the vehicle at any point in the trajectory, which is shown on the Primary Flight Display (PFD). It also provides caution and warning data and recommends the safest possible re-designated landing site when the nominal landing site is no longer within the capability of the vehicle. The PFD and the EMS allow the crew to manually fly an entry trajectory profile from entry interface until parachute deploy having the flexibility to manually steer the vehicle to a selected landing site that best satisfies the priorities of the crew. The entry display feeder provides data from the ENIS and other components of the GNC flight software to the displays at the proper rate and in the proper units. It also performs calculations that are specific to the entry displays and which are not made in any other component of the flight software. In some instances, it performs calculations identical to those performed by the onboard primary guidance algorithm to protect against a guidance system failure. These functions and the interactions between the entry display feeder and the other components of the EMS are described.

Baird, Darren↗
NASA NTRS1988

Tethered space recovery vehicle deployment/re-entry demonstration

The feasibility of utilizing existing Space Re-entry Vehicle (SRV) hardware for a Shuttle Orbiter-based tethered SRV deployment and re-entry demonstration using the Small Expendable Deployer System has been investigated. Options for mounting the SRV in the Orbiter, modifications and additions required to the existing SRV hardware have been defined. Flight demonstration scenarios from the Orbiter have been investigated, and re-entry motion and targeting uncertainties have been determined.

Florence, D.↗
NASA NTRS2013

Viability of 3 D Woven Carbon Cloth and Advanced Carbon-Carbon Ribs for Adaptive Deployable Entry Placement Technology (ADEPT) for Future NASA Missions

This paper describes aerothermodynamic and thermal structural testing that demonstrate the viability of three dimensional woven carbon cloth and advanced carbon-carbon (ACC) ribs for use in the Adaptive Deployable Entry Placement Technology (ADEPT). ADEPT is an umbrella-like entry system that is folded for stowage in the launch vehicle's shroud and deployed prior to reaching the atmeopheric interface. A key feature of the ADEPT concept is a lower ballistic coefficient for delivery of a given payload than seen with conventional, rigid body entry systems. The benefits that accrue from the lower ballistic coefficient incllude factor-of-ten reductions of deceleration forces and entry heating. The former enables consideration of new classes of scientific instruments for solar system exploration while the latter enables the design of a more efficient thermal protection system. The carbon cloth base lined for ADEPT has a dual use in that it serves as the thermal protection system and as the "skin" that transfers aerdynamic deceleration loads to its umbrella-like substructure. Arcjet testing described in this paper was conducted for some of the higher heating conditions for a future Venus mission using the ADEPT concept, thereby showing that the carbon cloth can perform in a relevant entry environment. Recently completed the thermal structural testing of the cloth attached to a representative ACC rib design is also described. Finally, this paper describes a preliminary engineering level code, based on the arcjet data, that can be used to estimate cloth thickness for future ADEPT missions and to predict carbon cloth performance in future arcjet tests.

ADEPT↗
NASA NTRS2013

Arcjet Testing of Woven Carbon Cloth for Use on Adaptive Deployable Entry Placement Technology

This paper describes arcjet testing and analysis that has successfully demonstrated the viability of three dimensional woven carbon cloth for dual use in the Adaptive Deployable Entry Placement Technology (ADEPT). ADEPT is an umbrella-like entry system that is folded for stowage in the launch vehicle s shroud and deployed in space prior to reaching the atmospheric interface. A key feature of the ADEPT concept is its lower ballistic coefficient for delivery of a given payload than those for conventional, rigid body entry systems. The benefits that accrue from the lower ballistic coefficient include factor of ten reductions of deceleration forces and entry heating. The former enables consideration of new classes of scientific instruments for solar system exploration while the latter enables the design of a more efficient thermal protection system. The carbon cloth now base lined for ADEPT has a dual use in that it serves as ADEPT s thermal protection system and as the "skin" that transfers aerodynamic deceleration loads to its umbrella-like substructure. The arcjet testing described in this paper was conducted for some of the higher heating conditions for a future Venus mission using the ADEPT concept, thereby showing that the carbon cloth can perform in a relevant entry environment. The ADEPT project considered the carbon cloth to be mission enabling and was carrying it as a major risk during Fiscal Year 2012. The testing and analysis reported here played a major role in retiring that risk and is highly significant to the success and possible adoption of ADEPT for future NASA missions. Finally, this paper also describes a preliminary engineering level code, based on the arcjet data, that can be used to estimate cloth thickness for future missions using ADEPT and to predict carbon cloth performance in future arcjet tests.

Arnold, James O.↗
NASA NTRS1985

Candidate techniques for non-intrusive measurement of flight boundary-layer properties

Techniques to measure gas composition and chemical activity in the hypersonic boundary layer associated with entry research vehicles and with a tethered satellite deployed at 100-125 km altitude are currently being investigated at the NASA-Langley Research Center. These measurements are appropriate for the study of the aerothermodynamic properties of these vehicles when operating with controlled trajectories in the thermosphere between 125 and 55 km. A number of the required quantitative physical measurements have been identified, and candidate techniques to obtain the measurements without perturbing the flow are being considered. Two candidate techniques are discussed in this paper: a mass spectrometer to analyze the gas composition at the vehicle surface, and Rayleigh scattering to obtain density, velocity, and temperature gradients from the surface to the edge of the boundary layer.

Wood, G. M., Jr.↗
NASA NTRS1988

A fault-tolerant avionics suite for an entry research vehicle

A highly-reliable avionics suite has been designed for an Entry Research Vehicle. The autonomous spacecraft would be deployed from the Space Shuttle Orbiter and perform a variety of aerodynamic and propulsive maneuvers which may be required for future space transportation system vehicles. The flight electronics consist of a central fault-tolerant processor, which is resilient to all first failures, reliably cross-strapped to redundant and distributed sets of sensors and effectors. This paper describes the preliminary design and analysis of the architecture which resulted from a fifteen month study by the Charles Stark Draper Laboratory for the NASA Langley Research Center. After a brief introduction to the design task, the architecture of the central flight computer and its interface to the vehicle are discussed. Following this, the method and results of the baseline reliability study for the avionic suite are presented.

Dzwonczyk, Mark↗
NASA NTRS2014

NASA Tech Briefs, March 2014

Topics include: Data Fusion for Global Estimation of Forest Characteristics From Sparse Lidar Data; Debris and Ice Mapping Analysis Tool - Database; Data Acquisition and Processing Software - DAPS; Metal-Assisted Fabrication of Biodegradable Porous Silicon Nanostructures; Post-Growth, In Situ Adhesion of Carbon Nanotubes to a Substrate for Robust CNT Cathodes; Integrated PEMFC Flow Field Design for Gravity-Independent Passive Water Removal; Thermal Mechanical Preparation of Glass Spheres; Mechanistic-Based Multiaxial-Stochastic-Strength Model for Transversely-Isotropic Brittle Materials; Methods for Mitigating Space Radiation Effects, Fault Detection and Correction, and Processing Sensor Data; Compact Ka-Band Antenna Feed with Double Circularly Polarized Capability; Dual-Leadframe Transient Liquid Phase Bonded Power Semiconductor Module Assembly and Bonding Process; Quad First Stage Processor: A Four-Channel Digitizer and Digital Beam-Forming Processor; Protective Sleeve for a Pyrotechnic Reefing Line Cutter; Metabolic Heat Regenerated Temperature Swing Adsorption; CubeSat Deployable Log Periodic Dipole Array; Re-entry Vehicle Shape for Enhanced Performance; NanoRacks-Scale MEMS Gas Chromatograph System; Variable Camber Aerodynamic Control Surfaces and Active Wing Shaping Control; Spacecraft Line-of-Sight Stabilization Using LWIR Earth Signature; Technique for Finding Retro-Reflectors in Flash LIDAR Imagery; Novel Hemispherical Dynamic Camera for EVAs; 360 deg Visual Detection and Object Tracking on an Autonomous Surface Vehicle; Simulation of Charge Carrier Mobility in Conducting Polymers; Observational Data Formatter Using CMOR for CMIP5; Propellant Loading Physics Model for Fault Detection Isolation and Recovery; Probabilistic Guidance for Swarms of Autonomous Agents; Reducing Drift in Stereo Visual Odometry; Future Air-Traffic Management Concepts Evaluation Tool; Examination and A Priori Analysis of a Direct Numerical Simulation Database for High-Pressure Turbulent Flows; and Resource-Constrained Application of Support Vector Machines to Imagery.

Source record↗
NASA NTRS1996

Sharp Hypervelocity Aerodynamic Research Probe

The objective of this flight demonstration is to deploy a slender-body hypervelocity aerodynamic research probe (SHARP) from an orbiting platform using a tether, deorbit and fly it along its aerothermal performance constraint, and recover it intact in mid-air. To accomplish this objective, two flight demonstrations are proposed. The first flight uses a blunt-body, tethered reentry experiment vehicle (TREV) to prove out tethered deployment technology for accurate entries, a complete SHARP electronics suite, and a new soft mid-air helicopter recovery technique. The second flight takes advantage of this launch and recovery capability to demonstrate revolutionary sharp body concepts for hypervelocity vehicles, enabled by new Ultra-High Temperature Ceramics (UHTCs) recently developed by Ames Research Center. Successful demonstration of sharp body hypersonic vehicle technologies could have radical impact on space flight capabilities, including: enabling global reentry cross range capability from Station, eliminating reentry communications blackout, and allowing new highly efficient launch systems incorporating air breathing propulsion and zeroth staging.

Bull, Jeffrey↗
NASA NTRS2012

Entry Atmospheric Flight Control Authority Impacts on GN and C and Trajectory Performance for Orion Exploration Flight Test 1

One of the key design objectives of NASA's Orion Exploration Flight Test 1 (EFT-1) is to execute a guided entry trajectory demonstrating GN&C capability. The focus of this paper is the ight control authority of the vehicle throughout the atmospheric entry ight to the target landing site and its impacts on GN&C, parachute deployment, and integrated performance. The vehicle's attitude control authority is obtained from thrusting 12 Re- action Control System (RCS) engines, with four engines to control yaw, four engines to control pitch, and four engines to control roll. The static and dynamic stability derivatives of the vehicle are determined to assess the inherent aerodynamic stability. The aerodynamic moments at various locations in the entry trajectory are calculated and compared to the available torque provided by the RCS system. Interaction between the vehicle's RCS engine plumes and the aerodynamic conditions are considered to assess thruster effectiveness. This document presents an assessment of Orion's ight control authority and its effectiveness in controlling the vehicle during critical events in the atmospheric entry trajectory.

McNamara, Luke W.↗
NASA NTRS1973

Advances in modeling aerodynamic decelerator dynamics.

The Viking entry vehicle uses a lines-first type of deployment in which the parachute, packed in a deployment bag, gets ejected rearward from the vehicle by a mortar. As the bag moves rearward, first the lines are unfurled and then the canopy. An analysis of the unfurling process is conducted, giving attention to longitudinal and rotational dynamics. It is shown that analytical modeling of aerodynamic systems provides significant information for a better understanding of the physics of the deployment process.

Whitlock, C. H.↗
NASA NTRS2005

Preliminary Report on Mission Design and Operations for Critical Events

Mission-critical events are defined in the Jet Propulsion Laboratory s Flight Project Practices as those sequences of events which must succeed in order to attain mission goals. These are dependent on the particular operational concept and design reference mission, and are especially important when committing to irreversible events. Critical events include main engine cutoff (MECO) after launch; engine cutoff or parachute deployment on entry, descent, and landing (EDL); orbital insertion; separation of payload from vehicle or separation of booster segments; maintenance of pointing accuracy for power and communication; and deployment of solar arrays and communication antennas. The purpose of this paper is to report on the current practices in handling mission-critical events in design and operations at major NASA spaceflight centers. The scope of this report includes NASA Johnson Space Center (JSC), NASA Goddard Space Flight Center (GSFC), and NASA Jet Propulsion Laboratory (JPL), with staff at each center consulted on their current practices, processes, and procedures.

Hayden, Sandra C.↗
NASA NTRS2009

Aerial Deployment and Inflation System for Mars Helium Balloons

A method is examined for safely deploying and inflating helium balloons for missions at Mars. The key for making it possible to deploy balloons that are light enough to be buoyant in the thin, Martian atmosphere is to mitigate the transient forces on the balloon that might tear it. A fully inflated Mars balloon has a diameter of 10 m, so it must be folded up for the trip to Mars, unfolded upon arrival, and then inflated with helium gas in the atmosphere. Safe entry into the Martian atmosphere requires the use of an aeroshell vehicle, which protects against severe heating and pressure loads associated with the hypersonic entry flight. Drag decelerates the aeroshell to supersonic speeds, then two parachutes deploy to slow the vehicle down to the needed safe speed of 25 to 35 m/s for balloon deployment. The parachute system descent dynamic pressure must be approximately 5 Pa or lower at an altitude of 4 km or more above the surface.

Lachenmeler, Tim↗