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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

Lessons Learned from the Investigation of an Anomalous Termination of BETTII

The Balloon Experimental Twin Telescope for Infrared Interferometry (BETTII) mission launched from Palestine, Texas in June 2017. After an exciting launch and successful cruise, the BETTII gondola suffered an anomalous event at termination. BETTII separated from its parachute and free-fell 136,000 feet into the west Texas desert. This event was classified as a "close-call" and investigated as such. We present here the recovery effort required to find the payload and extract the payload from its impact site. We also present lessons learned from the event and results from the investigation, the design for the next BETTII gondola, and a path forward for return to flight.

Veach, Todd J.↗
NASA NTRS2011

Monitoring Spacecraft Telemetry Via Optical or RF Link

A patent disclosure document discusses a photonic method for connecting a spacecraft with a launch vehicle upper-stage telemetry system as a means for monitoring a spacecraft fs health and status during and right after separation and deployment. This method also provides an efficient opto-coupled capability for prelaunch built-in-test (BIT) on the ground to enable more efficient and timely integration, preflight checkout, and a means to obviate any local EMI (electromagnetic interference) during integration and test. Additional utility can be envisioned for BIT on other platforms, such as the International Space Station (ISS). The photonic telemetry system implements an optical free-space link with a divergent laser transmitter beam spoiled over a significant cone angle to accommodate changes in spacecraft position without having to angle track it during deployment. Since the spacecraft may lose attitude control and tumble during deployment, the transmitted laser beam interrogates any one of several low-profile meso-scale retro-reflective spatial light modulators (SLMs) deployed over the surface of the spacecraft. The return signal beam, modulated by the SLMs, contains health, status, and attitude information received back at the launch vehicle. Very compact low-power opto-coupler technology already exists for the received signal (requiring relatively low bandwidths, e.g., .200 kbps) to enable transfer to a forward pass RF relay from the launch vehicle to TDRSS (Tracking and Data Relay Satellite System) or another recipient. The link would be active during separation and post-separation to monitor spacecraft health, status, attitude, or other data inventories until attitude recovery and ground control can be re-established. An optical link would not interfere with the existing upper stage telemetry and beacon systems, thus meeting launch vehicle EMI environmental constraints.

Fielhauer, K. B.↗
NASA NTRS1999

The Chandra X-Ray Observatory-Overview and Status

The Chandra X-Ray Observatory was launched early in the morning of 1999, July 23 by the Space Shuttle Columbia. The Shuttle launch was only the first step in placing NASA's latest great observatory into orbit. After release from the cargo bay, the Inertial Upper Stage performed two firings and separated from the observatory as planned. Finally, after five firings of Chandra's own Integral Propulsion System - the last of which took place 15 days after the initial launch - the observatory was placed in its highly elliptical orbit of 140,000 km apogee and 10,000 km perigee. After Observatory activation, the first x-rays focussed by the telescope were observed on 1999, August 12. Beginning with this initial observation one could conclude that the telescope had survived the launch environment and was operating as expected. The month following the opening of the sunshade door was spent adjusting the focus for each set of instrument configurations, determining the optical axis, calibrating the star camera, establishing the relative response functions, determining the energy scale(s), and performing a series of "publicity" images. Each observation proved to be far more revealing than was expected. Preliminary results will be presented and the status of the instrumentation on the observatory will be discussed.

Weisskopf, Martin C.↗
NASA NTRS2000

The Chandra X-Ray Observatory Overview

The Chandra X-Ray Observatory was launched early in the morning of July 23, 1999 by the Space Shuttle Columbia. The Shuttle launch was only the first step in placing NASA's latest great observatory into orbit. After release from the cargo bay, the Inertial Upper Stage performed two firings and separated from the observatory as planned. Finally, after five firings of Chandra's own Integral Propulsion System - the last of which took place 15 days after the initial launch - the observatory was placed in its highly elliptical orbit of 140,000 km apogee and 10,000 km perigee. After Observatory activation, the first x-rays focussed by the telescope were observed on August 12, 1999. Beginning with this initial observation one could conclude that the telescope had survived the launch environment and was operating as expected. The month following the opening of the sunshade door was spent adjusting the focus for each set of instrument configurations, determining the optical axis, calibrating the star camera, establishing the relative response functions, determining the energy scale(s), and performing a series of "publicity" images. Each observation proved to be far more revealing than was expected. Preliminary results will be presented and the status of the instrumentation on the observatory will be discussed.

Weisskopf, Martin C.↗
NASA NTRS2000

The Chandra X-Ray Observatory: An Overview

The Chandra X-Ray Observatory was launched early in the morning of July 23, 1999, by the Space Shuttle Columbia. The Shuttle launch was only the first step in placing NASA's latest great observatory into orbit. After release from the cargo bay, the Inertial Upper Stage performed two firings and separated from the observatory as planned. Finally, after five firings of Chandra's own Integral Propulsion System - the last of which took place 15 days after the initial launch - the observatory was placed in its highly elliptical orbit of 140,000 km apogee and 10,000 km perigee. After Observatory activation, the first x-rays focussed by the telescope were observed on 1999, August 12. Beginning with this initial observation one could conclude that the telescope had survived the launch environment and was operating as expected. The month following the opening of the sunshade door was spent adjusting the focus for each set of instrument configurations, determining the optical axis, calibrating the star camera, establishing the relative response functions, determining the energy scale(s), and performing a series of "publicity" images. Each observation proved to be far more revealing than was expected. Preliminary results will be presented and the status of the instrumentation on the observatory will be discussed.

Weisskopf, Martin C.↗
NASA NTRS2000

The Chandra X-Ray Observatory Overview

The Chandra X-Ray Observatory was launched early in the morning of July 23, 1999 by the Space Shuttle Columbia. The Shuttle launch was only the first step in placing NASA's latest great observatory into orbit. After release from the cargo bay, the Inertial Upper Stage performed two firings and separated from the observatory as planned. Finally, after five firings of Chandra's own Integral Propulsion System - the last of which took place 15 days after the initial launch - the observatory was placed in its highly elliptical orbit of 140,000 km apogee and 10,000 km perigee. After Observatory activation, the first x-rays focused by the telescope were observed on August 12, 1999. Beginning with this initial observation one could conclude that the telescope had survived the launch environment and was operating as expected. The month following the opening of the sunshade door was spent adjusting the focus for each set of instrument configurations, determining the optical axis, calibrating the star camera, establishing the relative response functions, determining the energy scale(s), and performing a series of "publicity" images. Each observation proved to be far more revealing than was expected. Preliminary results will be presented and the status of the instrumentation on the observatory will be discussed.

Weisskopf, Martin C.↗
NASA NTRS1999

Mars Orbit Rendezvous Strategy for the Mars 2003/2005 Sample Return Mission

The primary objective of the Mars 2003/2005 Sample Return Project is to return Martian surface materials to Earth from two different sites by the year 2008. The baseline mission plan relies heavily on the use of a Mars orbit rendezvous strategy similar to the lunar orbit rendezvous scheme used for the Apollo missions. The 2003 mission consists of a single spacecraft comprised of a Lander, Rover, and Mars ascent vehicle (MAV). The 2003 mission will be launched on a Delta-III-class launch vehicle in May/June 2003 and arrive at Mars in December 2003/January 2004. The Lander deploys the Rover to collect surface samples from several sites and return them to the Lander where they are transferred to a sample canister onboard the MAV. The MAV is launched into a low Mars orbit (targeted for 600 km circular, 45 deg inclination) and releases the sample canister to await retrieval by an Orbiter launched in 2005. (The sample canister is a passive vehicle with no maneuvering capability.) The duration of Mars surface operations is at most about 90 days. The 2005 mission consists of two separate spacecraft: a Lander/Rover/MAV spacecraft identical to that used for the 2003 mission and an Orbiter carrying an Earth Entry Vehicle (EEV). Both spacecraft will be launched on a single Ariane-5 in August 2005 and arrive at Mars in July/August 2006. A second sample canister is delivered to Mars orbit using the same scenario as was used for the 2003 mission. The Orbiter uses aerocapture for insertion into Mars orbit (targeted for 250 x 1400 km, 45 deg inclination). During its approximately one-year stay at Mars, the Orbiter will search for and attempt to rendezvous first with the 2003 sample canister and then with the 2005 sample canister. After retrieval, each sample canister is transferred to the EEV. The Orbiter departs Mars in July 2007 and returns to Earth in October 2008 on a trajectory targeted for landing at the Utah Test and Training Range (UTTR). After deploying the EEV, the Orbiter performs a deflection maneuver to avoid reentry into Earth's atmosphere.

DAmario, Louis A.↗
NASA NTRS2014

Testing Orions Fairing Separation System

Traditional fairing systems are designed to fully encapsulate and protect their payload from the harsh ascent environment including acoustic vibrations, aerodynamic forces and heating. The Orion fairing separation system performs this function and more by also sharing approximately half of the vehicle structural load during ascent. This load-share condition through launch and during jettison allows for a substantial increase in mass to orbit. A series of component-level development tests were completed to evaluate and characterize each component within Orion's unique fairing separation system. Two full-scale separation tests were performed to verify system-level functionality and provide verification data. This paper summarizes the fairing spring, Pyramidal Separation Mechanism and forward seal system component-level development tests, system-level separation tests, and lessons learned.

Martinez, Henry↗
NASA NTRS2013

Cradle-to-Grave Logistic Technologies for Exploration Missions

Human exploration missions under study are very limited by the launch mass capacity of exiting and planned vehicles. The logistical mass of crew items is typically considered separate from the vehicle structure, habitat outfitting, and life support systems. Consequently, crew item logistical mass is typically competing with vehicle systems for mass allocation. NASA is Advanced Exploration Systems (AES) Logistics Reduction and Repurposing (LRR) Project is developing four logistics technologies guided by a systems engineering cradle-to-grave approach to enable used crew items to augment vehicle systems. Specifically, AES LRR is investigating the direct reduction of clothing mass, the repurposing of logistical packaging, the processing of spent crew items to benefit radiation shielding and water recovery, and the conversion of trash to propulsion supply gases. The systematic implementation of these types of technologies will increase launch mass efficiency by enabling items to be used for secondary purposes and improve the habitability of the vehicle as the mission duration increases. This paper provides a description, benefits, and challenges of the four technologies under development and a status of progress at the mid ]point of the three year AES project.

Broyan, James L.↗
NASA NTRS2008

Ares V: Supporting Space Exploration from LEO to Beyond

This slide presentation reviews the development of the Ares V launch vehicle. The Ares V is designed as a heavy lift vehicle to send cargo only or in combination with the Ares I will send crews to the moon. The Ares V vehicle design is evolving and will make the authority to proceed milestone in 2011. The key milestones are given with the first mission flight in Fall of 2018. The Design Reference Mission (DRM) for the Ares V is reviewed, including a table with the events from liftoff through LSAM/CEV separation, in time and altitude. The Ares V will deliver 6 times more mass to orbit than currently launch vehicles can.

Sumrall, Phil↗
NASA NTRS2000

Vehicle Dynamics due to Magnetic Launch Propulsion

The field of Magnetic Levitation Lind Propulsion (MagLev) has been around for over 30 years, primarily in high-speed rail service. In recent years, however, NASA has been looking closely at MagLev as a possible first stage propulsion system for spacecraft. This approach creates a variety of new problems that don't currently exist with the present MagLev trains around the world. NASA requires that a spacecraft of approximately 120,000 lbs be accelerated at two times the acceleration of gravity (2g's). This produces a greater demand on power over the normal MagLev trains that accelerate at around 0.1g. To be able to store and distribute up to 3,000 Mega Joules of energy in less than 10 seconds is a technical challenge. Another problem never addressed by the train industry and, peculiar only to NASA, is the control of a lifting body through the acceleration of and separation from the MagLev track. Very little is understood about how a lifting body will react with external forces, Such as wind gusts and ground effects, while being propelled along on soft springs such as magnetic levitators. Much study needs to be done to determine spacecraft control requirements as well as what control mechanisms and aero-surfaces should be placed on the carrier. Once the spacecraft has been propelled down the track another significant event takes place, the separation of the spacecraft from the carrier. The dynamics involved for both the carrier and the spacecraft are complex and coupled. Analysis of the reaction of the carrier after losing, a majority of its mass must be performed to insure control of the carrier is maintained and a safe separation of the spacecraft is achieved. The spacecraft angle of attack required for lift and how it will affect the carriage just prior to separation, along with the impacts of around effect and aerodynamic forces at ground level must be modeled and analyzed to define requirements on the launch vehicle design. Mechanisms, which can withstand the expected forces and allow the spacecraft to separate at the desired time must be developed. Modeling and analysis of the full-scale system must be developed to derive the requirements for design of the separation mechanisms. A description of several MagLev demonstrators will be presented and the testing and analysis to model the demonstrators will be included.

Galaboff, Zachary J.↗
NASA NTRS2014

Logistics Reduction Technologies for Exploration Missions

Human exploration missions under study are very limited by the launch mass capacity of existing and planned vehicles. The logistical mass of crew items is typically considered separate from the vehicle structure, habitat outfitting, and life support systems. Consequently, crew item logistical mass is typically competing with vehicle systems for mass allocation. NASA's Advanced Exploration Systems (AES) Logistics Reduction and Repurposing (LRR) Project is developing five logistics technologies guided by a systems engineering cradle‐to‐grave approach to enable used crew items to augment vehicle systems. Specifically, AES LRR is investigating the direct reduction of clothing mass, the repurposing of logistical packaging, the use of autonomous logistics management technologies, the processing of spent crew items to benefit radiation shielding and water recovery, and the conversion of trash to propulsion gases. The systematic implementation of these types of technologies will increase launch mass efficiency by enabling items to be used for secondary purposes and improve the habitability of the vehicle as the mission duration increases. This paper provides a description and the challenges of the five technologies under development and the estimated overall mission benefits of each technology.

Broyan, James L., Jr.↗
NASA NTRS2009

Ares I-X: On the Threshold of Exploration

Ares I-X, the first flight of the Ares I crew launch vehicle, is less than a year from launch. Ares I-X will test the flight characteristics of Ares I from liftoff to first stage separation and recovery. The flight also will demonstrate the computer hardware and software (avionics) needed to control the vehicle; deploy the parachutes that allow the first stage booster to land in the ocean safely; measure and control how much the rocket rolls during flight; test and measure the effects of first stage separation; and develop and try out new ground handling and rocket stacking procedures in the Vehicle Assembly Building (VAB) and first stage recovery procedures at Kennedy Space Center (KSC) in Florida. All Ares I-X major elements have completed their critical design reviews, and are nearing final fabrication. The first stage--four-segment solid rocket booster from the Space Shuttle inventory--incorporates new simulated forward structures to match the Ares I five-segment booster. The upper stage, Orion crew module, and launch abort system will comprise simulator hardware that incorporates developmental flight instrumentation for essential data collection during the mission. The upper stage simulator consists of smaller cylindrical segments, which were transported to KSC in fall 2008. The crew module and launch abort system simulator were shipped in December 2008. The first stage hardware, active roll control system (RoCS), and avionics components will be delivered to KSC in 2009. This paper will provide detailed statuses of the Ares I-X hardware elements as NASA's Constellation Program prepares for this first flight of a new exploration era in the summer of 2009.

Davis, Stephan R.↗
NASA NTRS2010

A Dual Launch Robotic and Human Lunar Mission Architecture

This paper describes a comprehensive lunar exploration architecture developed by Marshall Space Flight Center's Advanced Concepts Office that features a science-based surface exploration strategy and a transportation architecture that uses two launches of a heavy lift launch vehicle to deliver human and robotic mission systems to the moon. The principal advantage of the dual launch lunar mission strategy is the reduced cost and risk resulting from the development of just one launch vehicle system. The dual launch lunar mission architecture may also enhance opportunities for commercial and international partnerships by using expendable launch vehicle services for robotic missions or development of surface exploration elements. Furthermore, this architecture is particularly suited to the integration of robotic and human exploration to maximize science return. For surface operations, an innovative dual-mode rover is presented that is capable of performing robotic science exploration as well as transporting human crew conducting surface exploration. The dual-mode rover can be deployed to the lunar surface to perform precursor science activities, collect samples, scout potential crew landing sites, and meet the crew at a designated landing site. With this approach, the crew is able to evaluate the robotically collected samples to select the best samples for return to Earth to maximize the scientific value. The rovers can continue robotic exploration after the crew leaves the lunar surface. The transportation system for the dual launch mission architecture uses a lunar-orbit-rendezvous strategy. Two heavy lift launch vehicles depart from Earth within a six hour period to transport the lunar lander and crew elements separately to lunar orbit. In lunar orbit, the crew transfer vehicle docks with the lander and the crew boards the lander for descent to the surface. After the surface mission, the crew returns to the orbiting transfer vehicle for the return to the Earth. This paper describes a complete transportation architecture including the analysis of transportation element options and sensitivities including: transportation element mass to surface landed mass; lander propellant options; and mission crew size. Based on this analysis, initial design concepts for the launch vehicle, crew module and lunar lander are presented. The paper also describes how the dual launch lunar mission architecture would fit into a more general overarching human space exploration philosophy that would allow expanded application of mission transportation elements for missions beyond the Earth-moon realm.

Jones, David L.↗
NASA NTRS1990

STS-36: Breakfast / Suit-Up / C-7 Ex / Launch and Landing at Edwards

Live footage shows the crew members of STS-36, Commander John O. Creighton, Pilot John H. Casper, and Mission Specialists Richard M. Mullane, David C. Hilmers, and Pierre J. Thuot, having the traditional breakfast, suiting up, and walking out to the Astro-Van. Scenes include panoramic views of the shuttle on the pad, main engine start, ignition, liftoff, and booster separation. The landing of Atlantis at Edwards Air Force Base is also seen. Several playback views from different cameras of both the launch and landing are also presented.

Source record↗
NASA NTRS1986

Interferometric phase velocity measurements in the auroral electrojet

A double-probe electric field detector and two spatially separated fixed-bias Langmuir probes were flown on a Taurus-Tomahawk sounding rocket launched from Poker Flat Research Range in March 1982. Interesting wave data have been obtained from about 10s of the downleg portion of the flight during which the rocket passed through the auroral electrojet. Here the electric field receiver and both density fluctuation (delta-n/n) receivers responded to a broad band of turbulence centered at 105 km-altitude and at frequencies generally below 4 kHz. Closer examination of the two delta-n/n turbulent waveforms reveals that they are correlated; from the phase difference between the two signals, the phase velocity of the waves in the rocket reference frame is inferred. The magnitude and direction of the observed phase velocity are consistent either with waves which travel at the ion sound speed or with waves which travel at the electron drift velocity. The observed phase velocity varies by about 50 percent over a 5 km altitude range, an effect which probably results from shear in the zonal neutral wind, although, unfortunately, no simultaneous neutral wind measurements exist to confirm this.

Labelle, J.↗
NASA NTRS1990

A transient response analysis of the space shuttle vehicle during liftoff

A proposed transient response method is formulated for the liftoff analysis of the space shuttle vehicles. It uses a power series approximation with unknown coefficients for the interface forces between the space shuttle and mobile launch platform. This allows the equation of motion of the two structures to be solved separately with the unknown coefficients at the end of each step. These coefficients are obtained by enforcing the interface compatibility conditions between the two structures. Once the unknown coefficients are determined, the total response is computed for that time step. The method is validated by a numerical example of a cantilevered beam and by the liftoff analysis of the space shuttle vehicles. The proposed method is compared to an iterative transient response analysis method used by Martin Marietta for their space shuttle liftoff analysis. It is shown that the proposed method uses less computer time than the iterative method and does not require as small a time step for integration. The space shuttle vehicle model is reduced using two different types of component mode synthesis (CMS) methods, the Lanczos method and the Craig and Bampton CMS method. By varying the cutoff frequency in the Craig and Bampton method it was shown that the space shuttle interface loads can be computed with reasonable accuracy. Both the Lanczos CMS method and Craig and Bampton CMS method give similar results. A substantial amount of computer time is saved using the Lanczos CMS method over that of the Craig and Bampton method. However, when trying to compute a large number of Lanczos vectors, input/output computer time increased and increased the overall computer time. The application of several liftoff release mechanisms that can be adapted to the proposed method are discussed.

Brunty, J. A.↗
NASA NTRS1995

GASCan 2 payload integration

This project, conducted in cooperation with the NASA Advanced Space Design Program, is part of an ongoing effort to place an experiment package into space. The goal of this project is to build and test flight-ready hardware that can be launched from the Space Shuttle. Get Away Special Canister 2 (GASCan 2) consists of three separate experiments. The Ionospheric Properties and Propagation Experiment (IPPE) determines effects of the ionosphere on radio wave propagation. The Microgravity Ignition experiment (MGI) tests the effects of combustion in a microgravity environment. The Rotational Fluid Flow experiment (RFF) examines fluid behavior under varying levels of gravity. This year the following tasks were completed: design of the IPPE antenna, X- and J-cell battery boxes, J-cell battery box enclosure, and structural bumpers; construction of the MGI canisters, MGI mounting brackets, IPPE antenna, and battery boxes; and the selection of the RFF's operating fluid and the analysis of the fluid behavior under microgravity test conditions.

Cody, Dennis J.↗