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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 451 records · Page 25
All sourcesNASA NTRSDOE OSTI
NASA NTRS2011

Reusable Solid Rocket Motor - Accomplishments, Lessons, and a Culture of Success

The Reusable Solid Rocket Motor represents the largest solid rocket motor ever flown and the only human rated solid motor. Each Reusable Solid Rocket Motor (RSRM) provides approximately 3-million lb of thrust to lift the integrated Space Shuttle vehicle from the launch pad. The motors burn out approximately 2 minutes later, separate from the vehicle and are recovered and refurbished. The size of the motor and the need for high reliability were challenges. Thrust shaping, via shaping of the propellant grain, was needed to limit structural loads during ascent. The motor design evolved through several block upgrades to increase performance and to increase safety and reliability. A major redesign occurred after STS-51L with the Redesigned Solid Rocket Motor. Significant improvements in the joint sealing systems were added. Design improvements continued throughout the Program via block changes with a number of innovations including development of low temperature o-ring materials and incorporation of a unique carbon fiber rope thermal barrier material. Recovery of the motors and post flight inspection improved understanding of hardware performance, and led to key design improvements. Because of the multidecade program duration material obsolescence was addressed, and requalification of materials and vendors was sometimes needed. Thermal protection systems and ablatives were used to protect the motor cases and nozzle structures. Significant understanding of design and manufacturing features of the ablatives was developed during the program resulting in optimization of design features and processing parameters. The project advanced technology in eliminating ozone-depleting materials in manufacturing processes and the development of an asbestos-free case insulation. Manufacturing processes for the large motor components were unique and safety in the manufacturing environment was a special concern. Transportation and handling approaches were also needed for the large hardware segments. The reusable solid rocket motor achieved significant reliability via process control, ground test programs, and postflight assessment. Process control is mandatory for a solid rocket motor as an acceptance test of the delivered product is not feasible. Process control included process failure modes and effects analysis, statistical process control, witness panels, and process product integrity audits. Material controls and inspections were maintained throughout the sub tier vendors. Material fingerprinting was employed to assess any drift in delivered material properties. The RSRM maintained both full scale and sub-scale test articles. These enabled continuous improvement of design and evaluation of process control and material behavior. Additionally RSRM reliability was achieved through attention to detail in post flight assessment to observe any shift in performance. The postflight analysis and inspections provided invaluable reliability data as it enables observation of actual flight performance, most of which would not be available if the motors were not recovered. These unique challenges, features of the reusable solid rocket motor, materials and manufacturing issues, and design improvements will be discussed in the paper.

Moore, Dennis R.↗
NASA NTRS2006

Inspection Of Spray On Foam Insulation (SOFI) Using and Microwave and Millimeter Wave Synthetic Aperture Focusing and Holography

The Space Shuttle Columbia's catastrophic failure is thought to have been caused by a dislodged piece of external tank spray on foam insulation (SOFI) striking the left wing of the orbiter causing significant damage to some of the reinforced carbon/carbon leading edge wing panels [1]. Microwave and millimeter wave nondestructive evaluation methods have shown great potential for inspecting SOFI for the purpose of detecting anomalies such as small air voids that may cause separation of the SOFI from the external tank during a launch [2,3]. These methods are capable of producing relatively high-resolution images of the interior of SOFI. Although effective, there are some advantages in using synthetic focusing methods as opposed to real focusing methods such as reduced probe size, the ability to determine depth from multiple views, and the ability to slice images due to sufficient range resolution. To this end, synthetic aperture focusing techniques (SAFT) were first pursued for this purpose and later wide-band microwave holography was implemented [4-7]. This paper presents the results of this investigation using frequency domain synthetic aperture focusing technique (FD-SAFT) and wide-band microwave holography methods illustrating their potential capabilities for inspecting the space shuttle's SOFI at millimeter wave frequencies.

Hepburn, F. L.↗
NASA NTRS2012

Solar Wind Charge Exchange and Local Hot Bubble X-Ray Emission with the DXL Sounding Rocket Experiment

The Diffuse X-ray emission from the Local Galaxy (DXL) sounding rocket is a NASA approved mission with a scheduled first launch in December 2012. Its goal is to identify and separate the X-ray emission of the SWCX from that of the Local Hot Bubble (LHB) to improve our understanding of both. To separate the SWCX contribution from the LHB. DXL will use the SWCX signature due to the helium focusing cone at 1=185 deg, b=-18 deg, DXL uses large area propostionai counters, with an area of 1.000 sq cm and grasp of about 10 sq cm sr both in the 1/4 and 3/4 keY bands. Thanks to the large grasp, DXL will achieve in a 5 minule flight what cannot be achieved by current and future X-ray satellites.

Galeazzi, M.↗
NASA NTRSDate not supplied

Hold Down and Release Mechanism Hardware Simulator for Ground Deployment Testing

Hold Down and Release Mechanisms (HDRMs) are a common piece of spacecraft hardware which restrain components during launch, transit, and later allow them to be extended, jettisoned, separated, etc. at the spacecraft’s destination. Due to prohibitively high cost, long lead times, and limited reset capability, testing of mechanisms that require HDRMs is often difficult. This project explored the possibility of designing HDRM-analogue devices that would replicate the function of flight HDRMs during ground-based testing for a fraction of the cost of a flight HDRM. These devices were designed with the requirements of the Spin Eject Mechanism to be used on the Mars Sample Return (MSR) Capture, Containment, and Return System (CCRS) in mind. A trade study was conducted to determine the optimal internal mechanism design of the device, comparing several mechanisms from existing HDRMs to several simpler, more novel mechanisms. CAD software was used to create a preliminary design of the device utilizing a mix of commercial off-the-shelf parts and custom-manufactured parts. Several straightforward analyses were conducted to determine the most effective models for the off-the-shelf parts. A prototype largely composed of 3D-printed components was also created to proof the design concept. Although this HDRM simulator was designed for a particular mechanism for the CCRS mission, its general design is useful for ground testing of a variety of mechanisms requiring similar HDRM simulators.

Kian Vilhauer↗
NASA NTRS1993

The initial flight anomalies of Skylab 1

At approximately 63 seconds into the flight of Skylab 1 on May 14, 1973, an anomaly occurred which resulted in the complete loss of the meteoroid shield around the orbital workshop. This was followed by the loss of one of the two solar array systems on the workshop and a failure of the inter stage adapter to separate from the S-II stage of the Saturn V launch vehicle. The investigation reported herein identified the most probable cause of this flight anomaly to be the breakup and loss of the meteoroid shield due to aerodynamic loads that were not accounted for in its design. The breakup of the meteoroid shield, in turn, broke the tie downs that secured one of the solar array systems to the workshop. Complete loss of this solar array system occurred at 593 seconds when the exhaust plume of the S-II stage retro-rockets impacted the partially deployed solar array system. Falling debris from the meteoroid shield also damaged the S-II inter stage adapter ordnance system in such a manner as to preclude separation. Of several possible failure modes of the meteoroid shield that were identified, the most probable in this particular flight was internal pressurization of its auxiliary tunnel which acted to force the forward end of the meteoroid shield away from the shell of the workshop and into the supersonic air stream. The pressurization of the auxiliary tunnel was due to the existence of several openings in the aft region of the tunnel. Another possible failure mode was the separation of the leading edge of the meteoroid shield from the shell of the workshop (particularly in the region of the folded ordnance panel) of sufficient extent to admit ram air pressures under the shield.

Source record↗
NASA NTRS2020

A Deformed Geometry Synthesis Technique for Determining Stacking and Cryogenically Induced Preloads for the Space Launch System

The Space Launch System (SLS) stacking and Core Stage (CS) fueling induce significant preloads that contribute to the liftoff pad separation “twang”. To accurately capture this, an approach is required that can replicate the physics of all SLS physical stacking steps, CS cryogenic shrinkage, associated geometric nonlinearities, and the transient behavior and decay of the preloads with changing boundary conditions as the vehicle separates from the pad. The Deformed Geometry Synthesis (DGS) approach presented here satisfies the above requirements. DGS determines induced preloads by modeling components in their deformed geometry states and then enforcing compatibility by closing the resulting “deadbands”. DGS seamlessly integrates into the multibody modal synthesis framework and does not require the use of artificial external loads to enforce preloads or post-processing steps to remove their influence. Since DGS iterates to solve for the deformed state inclusive of geometric nonlinearities, running linearized parametrics to exercise different potential orientations of ball jointed struts that connect the CS to Boosters for cryo-shrinkage analyses is entirely avoided. Relative to the transient behavior and decay of stacking and cryo-induced preloads with SLS liftoff pad separation, this is an area of considerable interest to the SLS program. To capture this in the most accurate way possible, DGS algorithms are designed to work with Henkel-Mar nonlinear pad separation algorithms which operate on the separating longitudinal and lateral degrees of freedom (DoFs) between the vehicle and the pad. As the separating DoFs release, in whatever manner as dictated by the interface geometries, interface loads and interface flexibilities as well as the external loading on the vehicle, the subject preloads generate a complex twang/decay time-trace as dictated by the physics of the problem. This paper presents DGS numerical verification against the closed-form solution for Timoshenko’s 3 ball-jointed strut preload problem. This problem is then extended by the authors to the geometric nonlinear case where DGS is compared to the Newton-Raphson solution of the nonlinear equations. Next, DGS is utilized to solve the SLS stacking and cryogenic shrinkage coupled loads analyses. Finally, Henkel-Mar pad separation simulations are executed that isolate the impact of the induced preloads’ twang and decay characteristics.

Space Launch System↗
NASA NTRS2010

An Up-Close Look at Ares I-X

On October 28, 2009, one day later than the originally planned launch date, the Ares I-X suborbital test flight roared into the Florida sky. Flying its preplanned parabolic arc over the Atlantic, the development test vehicle for the Ares I crew launch vehicle performed as advertised, executing a perfect liftoff, 90-degree roll maneuver, ascent, and separation before its upper and lower stages descended into the ocean 150 miles downrange. This test flight, while carrying no astronauts, marked a major milestone for NASA, which had not flown a test launch of a human-rated rocket since the first flight of the Space Shuttle in 1981. During the flight, over 700 sensors collected over 900 measurements, which NASA will apply to validating the engineering models they used to design the vehicle in the first place. That data, telemetered to the ground and stored in a flight recorder onboard, was the primary "payload" of the mission.

Leahy, Bart↗
NASA NTRS1961

Vanguard Satellite Separation Mechanisms

Early in the Vanguard program it became apparent that a thoroughly reliable means of separating the satellite packages from the third-stage rocket would be required. A completely self -contained standard mechanism was developed with redundant firing circuits for use on both test vehicles and satellite-launching vehicles. A change in the experimental objectives of the test-vehicle payload units necessitated modification of some of the standard separation mechanisms. A strap, pull-pin, girth-ring separation device was developed which employed the basic actuation of the standard mechanisms. Evidence of residual burning of the third stage made it necessary to delay separation longer than the time designed into the long-delay separation device. The standard separation mechanism was modified and integrated with the satellite command receiver system so that a ground command after third-stage burnout would cause separation. Flight performance of the various separation mechanisms proved their reliability; they performed without failure in all Vanguard launchings.

Baumann, Robert C.↗
NASA NTRSDate not supplied

SLS Block-1B USA Separation Analysis and Requirements Evaluation Using the CLVTOPS Toolchain

Ensuring that rocket stage separation events provide positive clearance is critical to avoid loss of mission or crew. NASA's Marshall Space Flight Center (MSFC) has developed a cutting-edge toolchain to address this type of problem and it was used to set abstracted impulse requirements on, and to analyze the results of, the in-space separation event of the Universal Stage Adapter (USA) and the Exploration Upper Stage (EUS) of NASA's Space Launch System (SLS) Block-1B configuration. The toolchain is used as a hardware simulation to confirm positive body-to-body clearance during the separation event. It is also used to create a requirements-space simulation, which helps inform requirements as the hardware design matures.

Zach Muscha↗
NASA NTRSDate not supplied

Near-Field Separation Analysis of SLS Block-1 and Block-1B Secondary Payloads

The NASA Space Launch System (SLS) Program Block-1 and Block-1B missions are expected to carry a number of CubeSat secondary payloads (SPL) to space, where they will be ejected from internal SLS structures to begin their own missions. This near-field ejection process must be closely evaluated to ensure that no SPL contacts the SLS structures, as contacts could result in SPL component damage, mission loss, or aberrant SPL trajectories. To this end, multiple SPL ejection events were analyzed for both Block-1 and Block-1B mission configurations using CLVTOPS, a NASA Marshall Space Flight Center (MSFC) developed multi-body dynamics, proximity analysis, and visualization toolchain. This paper details the SPL-to-SLS clearance assessment process, with a focus on simulation setup and results as well as SPL design requirement analysis. Relevant CLVTOPS, statistics, and mission backgrounds are covered; important constraints, difficulties, and assumptions are also documented. Furthermore, the importance of SPL housing geometry on near-field vehicle clearance is highlighted, and examples are shown.

Jared Rucker↗
NASA NTRS2008

Integrated Testing Approaches for the NASA Ares I Crew Launch Vehicle

The Ares I crew launch vehicle is being developed by the U.S. National Aeronautics and Space Administration (NASA) to provide crew and cargo access to the International Space Station (ISS) and, together with the Ares V cargo launch vehicle, serves as a critical component of NASA's future human exploration of the Moon. During the preliminary design phase, NASA defined and began implementing plans for integrated ground and flight testing necessary to achieve the first human launch of Ares I. The individual Ares I flight hardware elements - including the first stage five segment booster (FSB), upper stage, and J-2X upper stage engine - will undergo extensive development, qualification, and certification testing prior to flight. Key integrated system tests include the upper stage Main Propulsion Test Article (MPTA), acceptance tests of the integrated upper stage and upper stage engine assembly, a full-scale integrated vehicle ground vibration test (IVGVT), aerodynamic testing to characterize vehicle performance, and integrated testing of the avionics and software components. The Ares I-X development flight test will provide flight data to validate engineering models for aerodynamic performance, stage separation, structural dynamic performance, and control system functionality. The Ares I-Y flight test will validate ascent performance of the first stage, stage separation functionality, validate the ability of the upper stage to manage cryogenic propellants to achieve upper stage engine start conditions, and a high-altitude demonstration of the launch abort system (LAS) following stage separation. The Orion 1 flight test will be conducted as a full, un-crewed, operational flight test through the entire ascent flight profile prior to the first crewed launch.

Taylor, James L.↗
NASA NTRS2014

An Analysis and Review of Measures and Relationships in Space Transportation Affordability

The affordability of transportation to or from space is of continued interest across numerous and diverse stakeholders in our aerospace industry. Such an important metric as affordability deserves a clear understanding among stakeholders about what is meant by affordability, costs, and related terms, as otherwise it's difficult to see where specific improvements are needed or where to target specific investments. As captured in the famous words of Lewis Carroll, "If you don't know where you are going, any road will get you there". As important as understanding a metric may be, with terms such as costs, prices, specific costs, average costs, marginal costs, etc., it is equally important to understand the relationship among these measures. In turn, these measures intermingle with caveats and factors that introduce more measures in need of a common understanding among stakeholders. These factors include flight rates, capability, and payload. This paper seeks to review the costs of space transportation systems and the relationships among the many factors involved in costs from the points of view of diverse decision makers. A decision maker may have an interest in acquiring a single launch considering the best price (along with other factors in their business case), or an interest in many launches over time. Alternately, a decision maker may have a specific interest in developing a space transportation system that will offer certain prices, or flight rate capability, or both, at a certain up-front cost. The question arises for the later, to reuse or to expend? As it is necessary in thinking about the future to clearly understand the past and the present, this paper will present data and graphics to assist stakeholders in visualizing trends and the current state of affairs in the launch industry. At all times, raw data will be referenced (or made available separately) alongside detailed explanations about the data, so as to avoid the confusion or misleading conclusions that occur more often than not with complex graphs or statements when such context is lacking.

affordability↗
NASA NTRS1974

The development of sine vibration test requirements for Viking lander capsule components

In connection with the Viking project for exploring the planet Mars, two identical spacecraft, each consisting of an orbiter and a lander, will be launched in the third quarter of 1975. Upon arrival at the planet, the Viking lander will separate from the Viking orbiter and descend to a soft landing at a selected site on the Mars surface. It was decided to perform a sine vibration test on the Viking spacecraft, in its launch configuration, to qualify it for the booster-induced transient-dynamic environment. It is shown that component-level testing is a cost- and schedule-effective prerequisite to the system-level, sine-vibration test sequences.

Barrett, S.↗
NASA NTRS2007

Method for deploying multiple spacecraft

A method for deploying multiple spacecraft is disclosed. The method can be used in a situation where a first celestial body is being orbited by a second celestial body. The spacecraft are loaded onto a single spaceship that contains the multiple spacecraft and the spacecraft is launched from the second celestial body towards a third celestial body. The spacecraft are separated from each other while in route to the third celestial body. Each of the spacecraft is then subjected to the gravitational field of the third celestial body and each of the spacecraft assumes a different, independent orbit about the first celestial body. In those situations where the spacecraft are launched from Earth, the Sun can act as the first celestial body, the Earth can act as the second celestial body and the Moon can act as the third celestial body.

Sharer, Peter J.↗
NASA NTRS2014

Three Orbital Burns to Molniya Orbit via Shuttle Centaur G Upper Stage

An unclassified analytical trajectory design, performance, and mission study was done for the 1982-86 joint NASA-USAF Shuttle/Centaur G upper stage development program to send performance-demanding payloads to high orbits such as Molniya using an unconventional orbit transfer. This optimized three orbital burn transfer to Molniya orbit was compared to the then-baselined two burn transfer. The results of the three dimensional trajectory optimization performed include powered phase steering data and coast phase orbital element data. Time derivatives of the orbital elements as functions of thrust components were evaluated and used to explain the optimization's solution. Vehicle performance as a function of parking orbit inclination was given. Performance and orbital element data was provided for launch windows as functions of launch time. Ground track data was given for all burns and coasts including variation within the launch window. It was found that a Centaur with fully loaded propellant tanks could be flown from a 37deg inclination low Earth parking orbit and achieve Molniya orbit with comparable performance to the baselined transfer which started from a 57deg inclined orbit: 9,545 lb vs. 9,552 lb of separated spacecraft weight respectively. There was a significant reduction in the need for propellant launch time reserve for a one hour window: only 78 lb for the three burn transfer vs. 320 lb for the two burn transfer. Conversely, this also meant that longer launch windows over more orbital revolutions could be done for the same amount of propellant reserve. There was no practical difference in ground tracking station or airborne assets needed to secure telemetric data, even though the geometric locations of the burns varied considerably. There was a significant adverse increase in total mission elapsed time for the three vs. two burn transfer (12 vs. 11/4 hrs), but could be accommodated by modest modifications to Centaur systems. Future applications were discussed. The three burn transfer was found to be a viable, arguably preferable, alternative to the two burn transfer.

Orbital trajectory design↗
NASA NTRS1990

Aerodynamic separation and glideback of a Mach 3 staged booster

A study has been conducted of the staging maneuver for a two-stage, vertical-takeoff, launch vehicle and the subsequent glideback of the booster to a launch site runway. The booster and orbiter are both winged, fully reusable, and have liquid rocket main propulsion. for the staging maneuver, which nominally occurs at Mach 3, a flight control system and separation technique were devised which produced successful separations over a wide range of staging angles of attack. A guidance algorithm was developed for the unpowered glide-back of the booster from the completion of the staging maneuver to touchdown on a runway in the vicinity of the launch site. Using this guidance technique, the booster successfully touched down on the runway while being subjected to a wide range of seasonal and perturbed atmospheric conditions with and without wids as well as errors in staging state conditions and errors in predicted aerodynamics.

Naftel, J. C.↗
NASA NTRS2005

Nondestructive Evaluation of Space Shuttle External Tank Foam

During launch of Shuttle Columbia, mission STS-107, a large piece of spray on foam insulation (SOFI) separated from the external tank left bipod ramp area impacting the shuttle orbiter left wing leading edge. "Analysis showed that this large piece of foam struck Columbia on the underside of the left wing after launch. Later, analysis showed that the larger piece struck Columbia on the underside of the left wing, around Reinforced Carbon-Carbon (RCC) panels 5 through 9, at 81.9 seconds after launch. Further photographic analysis revealed that the large foam piece was approximately 21 to 27 inches long and 12 to 18 inches wide and was moving at a relative velocity to the Shuttle stack of 625 to 840 feet per second (416 to 573 miles per hour) at the time of impact." This impact damaged the wing leading edge resulting in loss of orbiter thermal protection. The piece of errant foam was part of a bipod ramp which was designed to meet thermal and aerodynamic requirements in that region of the external tank (ET).

Thom, Robert L.↗
NASA NTRS2015

Three Orbital Burns to Molniya Orbit Via Shuttle_Centaur G Upper Stage

An unclassified analytical trajectory design, performance, and mission study was done for the 1982 to 1986 joint National Aeronautics and Space Administration (NASA)-United States Air Force (USAF) Shuttle/Centaur G upper stage development program to send performance-demanding payloads to high orbits such as Molniya using an unconventional orbit transfer. This optimized three orbital burn transfer to Molniya orbit was compared to the then-baselined two burn transfer. The results of the three dimensional trajectory optimization performed include powered phase steering data and coast phase orbital element data. Time derivatives of the orbital elements as functions of thrust components were evaluated and used to explain the optimization's solution. Vehicle performance as a function of parking orbit inclination was given. Performance and orbital element data was provided for launch windows as functions of launch time. Ground track data was given for all burns and coasts including variation within the launch window. It was found that a Centaur with fully loaded propellant tanks could be flown from a 37 deg inclination low Earth parking orbit and achieve Molniya orbit with comparable performance to the baselined transfer which started from a 57 deg inclined orbit: 9,545 versus 9,552 lb of separated spacecraft weight, respectively. There was a significant reduction in the need for propellant launch time reserve for a 1 hr window: only 78 lb for the three burn transfer versus 320 lb for the two burn transfer. Conversely, this also meant that longer launch windows over more orbital revolutions could be done for the same amount of propellant reserve. There was no practical difference in ground tracking station or airborne assets needed to secure telemetric data, even though the geometric locations of the burns varied considerably. There was a significant adverse increase in total mission elapsed time for the three versus two burn transfer (12 vs. 1-1/4 hr), but could be accommodated by modest modifications to Centaur systems. Future applications were discussed. The three burn transfer was found to be a viable, arguably preferable, alternative to the two burn transfer.

Orbital trajectory design↗