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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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At least 253 records · Page 14
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NASA NTRS2005

Low-Cost Innovation in Spaceflight: The Near Earth Asteroid Rendezvous (NEAR) Shoemaker Mission

On a spring day in 1996, at their research center in the Maryland countryside, representatives from the Johns Hopkins University Applied Physics Laboratory (APL) presented Administrator Daniel S. Goldin of the National Aeronautics and Space Administration (NASA) with a check for $3.6 million. 1 Two and a half years earlier, APL officials had agreed to develop a spacecraft capable of conducting an asteroid rendezvous and to do so for slightly more than $122 million. This was a remarkably low sum for a spacecraft due to conduct a planetaryclass mission. By contrast, the Mars Observer spacecraft launched in 1992 for an orbital rendezvous with the red planet had cost $479 million to develop, while the upcoming Cassini mission to Saturn required a spacecraft whose total cost was approaching $1.4 billion. In an Agency accustomed to cost overruns on major missions, the promise to build a planetary-class spacecraft for about $100 million seemed excessively optimistic.

McCurdy, Howard E.↗
NASA NTRS1987

Ka-band (32 GHz) spacecraft development plan

A road map for the development of a protoflight 32 GHz spacecraft solid state transmitter is given. The major milestones include the development of device and component technology required for use in the spaceborne experimental and operational transmitter systems. Two experimental spacecraft transmitter systems are envisioned: first, a low power beacon, to determine the performance of a 32 GHz downlink communications system; and second, an array feed, to further verify the results of the first experiment and service as a test bed for technology required for an operational system. The first experiment was proposed to NASA Headquarters for flight aboard the Mars Observer spacecraft with spacecraft integration in early 1989. The second is to be available for integration aboard a spacecraft such as the Comet Rendezvous Asteroid Flyby (CRAF) mission in the 1990 time frame. These experimental systems are to lead to the development of a protoflight transmitter for subsequent spacecraft integration in 1992, the time frame of the Cassini mission to Saturn.

Riley, A. L.↗
NASA NTRSDate not supplied

Integrated Design Results for the MSR SRC Mars Ascent Vehicle

The primary mission of the NASA Mars Sample Return (MSR) Campaign is to return samples of the Martian surface to Earth for scientific study. As part of this campaign, NASA is developing a Mars Ascent Vehicle (MAV). This vehicle must survive an approximate two year journey to the Martian surface as a payload aboard a separate lander spacecraft. After residing on the surface for another year, the MAV will carry a payload of samples into orbit. From there, following ejection from the MAV, the samples will rendezvous with an Earth return spacecraft for capture, and ultimately, return to Earth.The design of the MAV represents a number of unique challenges, as no launch vehicle has ever left the surface of a planet other than Earth. Although conceptual designs for a MAV have been in various levels of development since the 1970s, none have achieved the level of fidelity and support that exists in the current MSR-MAV design. Early MSR-MAV concept studies examined multiple methods of propulsion, ultimately deciding that a Two Stage to Orbit (TSTO) solid propulsion vehicle would provide the most capable performance in a Martian environment. Following this key architecture decision, the vehicle design was further matured to a Solid-Solid Guided-Guided (SSGG) architecture for NASA Key Decision Point A (KDP-A). Although the SSGG design was able to meet all mission constraints, concerns were raised regarding limited mass margin on other elements of the MSR campaign at such an early phase. A design challenge was issued to reduce MAV total mass by as much as possible. It was ultimately determined that by moving a number of components of the vehicle second stage to the first stage, the overall vehicle mass could be reduced significantly. The new design featured a much smaller and completely unguided second stage. This paper describes the resultant Solid-Solid Guided-Unguided (SSGU) MAV design concept developed as part of the Systems Requirement Cycle (SRC). This design was developed primarily by NASA Marshall Space Flight Center (MSFC), in association with NASA Jet Propulsion Laboratory (JPL) and NASA Langley Research Center (LaRC). The TSTO vehicle includes one solid rocket motor per stage. As the vehicle second stage is unguided, it features spin-stabilization to maintain vehicle stability during flight. An electromechanically actuated Thrust Vector Control (TVC) and a monopropellant Reaction Control System (RCS) is employed for active guidance on the first stage. The vehicle is designed to deliver up to 0.47kg of Martian samples to a Mars circular orbit of 380km at 27° inclination. Due to the extremely unique design constraints of this mission, and a recent transition to a Risk Class A posture, the MAV team was compelled to devise unconventional solutions to the vehicle design. The detailed design and analysis of these subsystems and the vehicle as a whole are discussed in this paper relative to all of the engineering disciplines involved.

MSR↗
NASA NTRSDate not supplied

Integrated Design Results for the MSR SRC Mars Ascent Vehicle

The primary mission of the NASA Mars Sample Return (MSR) Campaign is to return samples of the Martian surface to Earth for scientific study. As part of this campaign, NASA is developing a Mars Ascent Vehicle (MAV). This vehicle must survive an approximate two-year journey to the Martian surface as a payload aboard a separate lander spacecraft. After residing on the surface for another year, the MAV will carry a payload of samples into orbit. From there, following ejection from the MAV, the samples will rendezvous with an Earth return spacecraft for capture, and ultimately, return to Earth. The design of the MAV represents a number of unique challenges, as no launch vehicle has ever left the surface of a planet other than Earth. Although conceptual designs for a MAV have been in various levels of development since the 1970s, none have achieved the level of fidelity and support that exists in the current MSR-MAV design. Early MSR-MAV concept studies examined multiple methods of propulsion, ultimately deciding that a Two-Stage-to-Orbit (TSTO) solid propulsion vehicle would provide the most capable performance in a Martian environment. Following this key architecture decision, the vehicle design was further matured to a Solid-Solid Guided-Guided (SSGG) architecture for NASA Key Decision Point A (KDP-A). Although the SSGG design was able to meet all mission constraints, concerns were raised regarding limited mass margin on other elements of the MSR campaign at such an early phase. A design challenge was issued to reduce MAV total mass by as much as possible. It was ultimately determined that by moving a number of components of the vehicle second stage to the first stage, the overall vehicle mass could be reduced significantly. The new design features a much smaller and completely unguided second stage. This paper describes the resultant Solid-Solid Guided-Unguided (SSGU) MAV design concept developed as part of the Systems Requirement Cycle (SRC). This design was developed primarily by NASA Marshall Space Flight Center (MSFC), in association with NASA Jet Propulsion Laboratory (JPL) and NASA Langley Research Center (LaRC). The TSTO vehicle includes one solid rocket motor per stage. As the vehicle second stage is unguided, it features spin-stabilization to maintain vehicle stability during flight. An electromechanically actuated Thrust Vector Control (TVC) system and a monopropellant Reaction Control System (RCS) are employed for active guidance on the first stage. The vehicle is designed to deliver up to 0.47kg of Martian samples to a Mars circular orbit of 380km at 27° inclination. Due to the extremely unique design constraints of this mission, and a recent transition to a Risk Class A posture, the MAV team was compelled to devise unconventional solutions to the vehicle design. The detailed design and analysis of these subsystems and the vehicle as a whole are discussed in this paper relative to all of the engineering disciplines involved.

MAV↗
NASA NTRS1987

Comet Rendezvous Asteroid Flyby - First Mariner Mark II

The Mariner Mark II, a novel, low-cost series of spacecraft, will find its initial use during the Comet Rendezvous Asteroid Flyby (CRAF) Project; the spacecraft is expected to obtain the first observations of asteroids, and to return over three years' worth of detailed science data from the vicinity of Comet Tempel 2. The second Mariner Mark II spacecraft, designated 'Cassini' will incorporate an ESA Titan probe. Attention is given to a possible solar/electric earth gravity-assist trajectory for a mission to Tempel 2.

Draper, Ron↗
NASA NTRS2012

Pose Measurement Performance of the Argon Relative Navigation Sensor Suite in Simulated Flight Conditions

Argon is a flight-ready sensor suite with two visual cameras, a flash LIDAR, an on- board flight computer, and associated electronics. Argon was designed to provide sensing capabilities for relative navigation during proximity, rendezvous, and docking operations between spacecraft. A rigorous ground test campaign assessed the performance capability of the Argon navigation suite to measure the relative pose of high-fidelity satellite mock-ups during a variety of simulated rendezvous and proximity maneuvers facilitated by robot manipulators in a variety of lighting conditions representative of the orbital environment. A brief description of the Argon suite and test setup are given as well as an analysis of the performance of the system in simulated proximity and rendezvous operations.

Galante, Joseph M.↗
NASA NTRS1995

JPL's Galileo Spacecraft Will Shoot an Interplanetary 'Bulls's Eye'

JPL's Galileo spacecraft, with a December 1995 rendezvous with Jupiter, shot an interplanetary 'bull's eye' when it released its probe to conduct history's first sampling of Jupiter's atmosphere. The probe will slam into Jupiter's hydrogen-helium atmosphere at 100,000 mph as the spacecraft flies overhead and receives and records the probe's data on the structure and composition of the atmosphere.

Galileo Jupiter Interplanetary Spacecraft↗
NASA NTRS2007

Development of a Deployable Nonmetallic Boom for Reconfigurable Systems of Small Modular Spacecraft

Launch vehicle payload capacity and the launch environment represent two of the most operationally limiting constraints on space system mass, volume, and configuration. Large-scale space science and power platforms as well as transit vehicles have been proposed that greatly exceed single-launch capabilities. Reconfigurable systems launched as multiple small modular spacecraft with the ability to rendezvous, approach, mate, and conduct coordinated operations have the potential to make these designs feasible. A key characteristic of these proposed systems is their ability to assemble into desired geometric (spatial) configurations. While flexible and sparse formations may be realized by groups of spacecraft flying in close proximity, flyers physically connected by active structural elements could continuously exchange power, fluids, and heat (via fluids). Configurations of small modular spacecraft temporarily linked together could be sustained as long as needed with minimal propellant use and reconfigured as often as needed over extended missions with changing requirements. For example, these vehicles could operate in extremely compact configurations during boost phases of a mission and then redeploy to generate power or communicate while coasting and upon reaching orbit. In 2005, NASA funded Phase 1 of a program called Modular Reconfigurable High-Energy Technology Demonstrator Assembly Testbed (MRHE) to investigate reconfigurable systems of small spacecraft. The MRHE team was led by NASA's Marshall Space Flight Center and included Lockheed Martin's Advanced Technology Center (ATC) in Palo Alto and its subcontractor, ATK. One of the goals of Phase 1 was to develop an MRHE concept demonstration in a relevant 1-g environment to highlight a number of requisite technologies. In Phase 1 of the MRHE program, Lockheed Martin devised and conducted an automated space system assembly demonstration featuring multipurpose free-floating robots representing Spacecraft in the newly built Controls and Automation Laboratory (CAL) at the ATC. The CAL lab features a 12' x 24' granite air-bearing table and an overhead simulated starfield. Among the technologies needed for the concept demo were mating interfaces allowing the spacecraft to dock and deployable structures allowing for adjustable separation between spacecraft after a rigid connection had been established. The decision to use a nonmetallic deployable boom for this purpose was driven by the MRHE concept demo requirements reproduced in Table 1.

Rehnmark, Fredrik↗
NASA NTRS1974

Navigation and performance analysis of a Mars surface sample return /MSSR/ mission

A Viking-derived MSSR mission involves a rendezvous in Mars orbit between the sample-carrying Mars Ascent Vehicle (MAV) and a modified Viking Orbiter spacecraft. Conditions for the automatic terminal rendezvous phase are established by an initial earth-controlled phase which uses very accurate relative state information obtained by multi-vehicle interferometric tracking data. A navigation analysis has demonstrated sufficient performance and control capability to satisfy those conditions prior to terminal rendezvous. Specific performance constraints for the mission are developed, leading to a sizing of vehicle weight requirements and trades applicable to a 1981 MSSR, launched by a single Titan IIIE/Centaur.

Satin, A. L.↗
NASA NTRS1962

RENDEZVOUS AND DOCKING TECHNIQUES

To implement the primary space mission of this decade – manned lunar exploration -- the operational assistance of rendezvous and docking is being considered as an alternative to possible problems in obtaining a boost vehicle capable of direct flight. This paper concentrates on the mechanization of the rendezvous and docking phase of such a lunar mission. As indicated in previous papers, rendezvous can take place in either an earth or lunar orbit (or on the lunar surface); it can involve the mating of stages, transfer of fuel, or the return of a shuttle to a “mother” ship; direct ascent or parking orbits can be used; the orbits can be circular or more general ellipses; finally, either or both of the spacecraft can participate in the rendezvous maneuvers. Since there is no intent here to discuss the pros and cons of each approach or to cover all possible mission profiles, guidance schemes and hardware configurations, one particular profile has been selected to display the significant features of most rendezvous missions. In many respects, rendezvous is less difficult than aircraft interception since the target is friendly and there is no severe time constraint. The latter feature allows considerable freedom of design especially with the great versatility of a human in the loop. This remains true even for a purely automatic mode. Perhaps the largest design problem concerns the selection of the mission to be implemented followed by the optimization or systems engineering of a mechanization from the multitude of possible schemes and techniques. A most important factor in the optimization is that of reliability and the redundancy, alternate or backup modes, etc., associated with the approach. While a primary system can be rather easily mechanized with modest equipment requirements, reliability considerations will result in additional hardware, tighter specifications and more safety factors (e. g., propellant margin). This is a very complex area and will only briefly be mentioned in the following discussion.

Rendezvous↗
NASA NTRS2013

A Survey of LIDAR Technology and Its Use in Spacecraft Relative Navigation

This paper provides a survey of modern LIght Detection And Ranging (LIDAR) sensors from a perspective of how they can be used for spacecraft relative navigation. In addition to LIDAR technology commonly used in space applications today (e.g. scanning, flash), this paper reviews emerging LIDAR technologies gaining traction in other non-aerospace fields. The discussion will include an overview of sensor operating principles and specific pros/cons for each type of LIDAR. This paper provides a comprehensive review of LIDAR technology as applied specifically to spacecraft relative navigation. HE problem of orbital rendezvous and docking has been a consistent challenge for complex space missions since before the Gemini 8 spacecraft performed the first successful on-orbit docking of two spacecraft in 1966. Over the years, a great deal of effort has been devoted to advancing technology associated with all aspects of the rendezvous, proximity operations, and docking (RPOD) flight phase. After years of perfecting the art of crewed rendezvous with the Gemini, Apollo, and Space Shuttle programs, NASA began investigating the problem of autonomous rendezvous and docking (AR&D) to support a host of different mission applications. Some of these applications include autonomous resupply of the International Space Station (ISS), robotic servicing/refueling of existing orbital assets, and on-orbit assembly.1 The push towards a robust AR&D capability has led to an intensified interest in a number of different sensors capable of providing insight into the relative state of two spacecraft. The present work focuses on exploring the state-of-the-art in one of these sensors - LIght Detection And Ranging (LIDAR) sensors. It should be noted that the military community frequently uses the acronym LADAR (LAser Detection And Ranging) to refer to what this paper calls LIDARs. A LIDAR is an active remote sensing device that is typically used in space applications to obtain the range to one or more points on a target spacecraft. As the name suggests, LIDAR sensors use light (typically a laser) to illuminate the target and measure the time it takes for the emitted signal to return to the sensor. Because the light must travel from the source, to

Christian, John A.↗
NASA NTRS1988

Determination of the Venus flyby orbits of the Soviet Vega probes using VLBI techniques

In December 1984, the Soviet Union launched two identical Vega spacecraft with the dual objective of exploring Venus and continuing to rendezvous with the comet Halley. The two Vega spacecraft encountered Venus in mid-June 1985 and successfully deployed entry probes and wind-measuring balloons into the Venus atmosphere. An objective of the Venus Balloon experiment was to measure the Venus winds using differential VLBI from the balloon and the flyby bus. NASA's Deep Space 64 meter subnet was part of a world wide network organized to collect data from the Vega probes and balloons. A critical element of this experiment was an accurate determination of the Venus relative flyby orbits of the Vega spacecraft during the 46 hour balloon lifetime. Venus flyby solutions were independently determined by the Soviets using two-way range and Doppler from Soviet stations and by JPL using one-way Doppler and VLBI data collected from the DSN. The Vega flyby solutions determined by the Soviets using a sparse two-way tracking strategy with JPL solutions using the DSN VLBI data to complement the Soviet data and with solutions using only one-way data collected by the DSN were compared.

Ellis, J.↗
NASA NTRS2012

Next Gen NEAR: Near Earth Asteroid Human Robotic Precursor Mission Concept

Asteroids have long held the attention of the planetary science community. In particular, asteroids that evolve into orbits near that of Earth, called near-Earth objects (NEO), are of high interest as potential targets for exploration due to the relative ease (in terms of delta V) to reach them. NASA's Flexible Path calls for missions and experiments to be conducted as intermediate steps towards the eventual goal of human exploration of Mars; piloted missions to NEOs are such example. A human NEO mission is a valuable exploratory step beyond the Earth-Moon system enhancing capabilities that surpass our current experience, while also developing infrastructure for future mars exploration capabilities. To prepare for a human rendezvous with an NEO, NASA is interested in pursuing a responsible program of robotic NEO precursor missions. Next Gen NEAR is such a mission, building on the NEAR Shoemaker mission experience at the JHU/APL Space Department, to provide an affordable, low risk solution with quick data return. Next Gen NEAR proposes to make measurements needed for human exploration to asteroids: to demonstrate proximity operations, to quantify hazards for human exploration and to characterize an environment at a near-Earth asteroid representative of those that may be future human destinations. The Johns Hopkins University Applied Physics Laboratory has demonstrated exploration-driven mission feasibility by developing a versatile spacecraft design concept using conventional technologies that satisfies a set of science, exploration and mission objectives defined by a concept development team in the summer of 2010. We will describe the mission concept and spacecraft architecture in detail. Configuration options were compared with the mission goals and objectives in order to select the spacecraft design concept that provides the lowest cost, lowest implementation risk, simplest operation and the most benefit for the mission implementation. The Next Gen NEAR spacecraft was designed to support rendezvous with a range of candidate asteroid targets and could easily be launched with one of several NASA launch vehicles. The Falcon 9 launch vehicle supports a Next Gen NEAR launch to target many near-Earth asteroids under consideration that could be reached with a C3 of 18 km2/sec2 or less, and the Atlas V-401 provides added capability supporting launch to NEAs that require more lift capacity while at the same time providing such excess lift capability that another payload of opportunity could be launch in conjunction with Next Gen NEAR. Next Gen NEAR will measure and interact with the target surface in ways never undertaken at an asteroid, and will prepare for first human precursor mission by demonstrating exploration science operations at an accessible NEO. This flexible mission and spacecraft design concept supports target selection based on upcoming Earth-based observations and also provides opportunities for co-manifest & international partnerships. JHU/APL has demonstrated low cost, low risk, high impact missions and this mission will help to prepare NASA for human NEO exploration by combining the best of NASA s human and robotic exploration capabilities.

Rivkin, Andrew S.↗
NASA NTRS2003

Demonstration of Autonomous Rendezvous Technology (DART) Project Summary

Since the 1960's, NASA has performed numerous rendezvous and docking missions. The common element of all US rendezvous and docking is that the spacecraft has always been piloted by astronauts. Only the Russian Space Program has developed and demonstrated an autonomous capability. The Demonstration of Autonomous Rendezvous Technology (DART) project currently funded under NASA's Space Launch Initiative (SLI) Cycle I, provides a key step in establishing an autonomous rendezvous capability for the United States. DART's objective is to demonstrate, in space, the hardware and software necessary for autonomous rendezvous. Orbital Sciences Corporation intends to integrate an Advanced Video Guidance Sensor and Autonomous Rendezvous and Proximity Operations algorithms into a Pegasus upper stage in order to demonstrate the capability to autonomously rendezvous with a target currently in orbit. The DART mission will occur in April 2004. The launch site will be Vandenburg AFB and the launch vehicle will be a Pegasus XL equipped with a Hydrazine Auxiliary Propulsion System 4th stage. All mission objectives will be completed within a 24 hour period. The paper provides a summary of mission objectives, mission overview and a discussion on the design features of the chase and target vehicles.

TImothy E Rumford↗
NASA NTRS1978

SEP ENCKE-87 and Halley rendezvous studies and improved S/C model implementation in HILTOP

Studies were conducted to determine the performance requirements for projected state-of-the-art SEP spacecrafts boosted by the Shuttle/IUS to perform a rendezvous with the comet Halley and a rendezvous with the comet Encke during its 1977 apparition. The spacecraft model of the standard HILTOP computer program was assumed. Numerical and graphical results summarizing the studies are presented.

Horsewood, J. L.↗
NASA NTRS1977

Solar electric propulsion thruster interactions with solar arrays

The effect of interactions of spacecraft-generated and naturally occurring plasmas with high voltage solar array components on an advanced solar electric propulsion system proposed for the Halley's Comet rendezvous mission was investigated. The spacecraft-generated plasma consists of mercury ions and neutralizing electrons resulting from the operation of ion thrusters (the charge-exchange plasma) and associated hollow cathode neutralizers. Quantitative results are given for the parasitic currents and power coupled into solar arrays with voltage fixed as a function of position on the array.

Parks, D. E.↗