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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 NTRS2006

Ares 1 First Stage Design, Development, Test, and Evaluation

The Ares I Crew Launch Vehicle (CLV) is an integral part of NASA s exploration architecture that will provide crew and cargo access to the International Space Station as well as low earth orbit support for lunar missions. Currently in the system definition phase, the CLV is planned to replace the Space Shuttle for crew transport in the post 2010 time frame. It is comprised of a solid rocket booster (SRB) first stage derived from the current Space Shuttle SRB, a liquid oxygen/hydrogen fueled second stage utilizing a derivative of the Apollo upper stage engine for propulsion, and a Crew Exploration Vehicle (CEV) composed of command and service modules. This paper deals with current design, development, test, and evaluation planning for the CLV first stage SRB. Described are the current overall point-of-departure design and booster subsystems, systems engineering approach, and milestone schedule requirements.

Williams, Tom↗
NASA NTRS2007

Staging Airliner Service

There is a general consensus building that historically high fuel prices and greater public awareness of the emissions that result from burning fuel are going to be long-term concerns for those who design, build, and operate airliners. The possibility of saving both fuel and reducing emissions has rekindled interest in breaking very long-range airline flights into multiple stages or even adopting in-flight refueling. It is likely that staging will result in lower fuel burn, and recent published reports have suggested that the savings are substantial, particularly if the airliner is designed from the outset for this kind of operation. Given that staging runs against the design and operation historical trend, this result begs for further attention. This paper will examine the staging question, examining both analytic and numeric performance estimation methodologies to quantify the likely amount of fuel savings that can be expected and the resulting design impacts on the airliner.

Hahn, Andrew S.↗
NASA NTRS2008

Two-Stage Passive Vibration Isolator

The design and testing of a structural system were implemented to hold the optics of the planned Space Interferometry Mission (SIM) at positions and orientations characterized by vibrational translation and rotation errors of no more than a few nanometers or a few milliarcseconds, respectively. Much of the effort was devoted to a test bed for verifying the predicted behavior of a vibration- isolation structural subsystem working together with an active control system for positioning and orienting the SIM optics. There was considerable emphasis on the vibration-isolation subsystem, which was passive and comprised two stages. The main sources of vibration were six reaction wheels in an assembly denoted the "backpack." The first vibration-isolation stage consisted of hexapod isolator mounts - one for each reaction wheel - characterized by a natural vibration frequency of 10 Hz. The second stage was a set of three beams, disposed between the backpack and the structure that held the SIM optics, that were flexured such that they transmitted only bending loads, with a natural vibrational frequency and damping of about 5 Hz and 4 percent, respectively. Preliminary test results were presented and characterized as demonstrating the effectiveness of the two-stage vibration-isolation design.

Goullioud, Renaud↗
NASA NTRS2008

Initial Assessment of the Ares I-X Launch Vehicle Upper Stage to Vibroacoustic Flight Environments

The Ares I launch vehicle will be NASA s first new launch vehicle since 1981. Currently in design, it will replace the Space Shuttle in taking astronauts to the International Space Station, and will eventually play a major role in humankind s return to the Moon and eventually to Mars. Prior to any manned flight of this vehicle, unmanned test readiness flights will be flown. The first of these readiness flights, named Ares I-X, is scheduled to be launched in April 2009. The NASA Glenn Research Center is responsible for the design, manufacture, test and analysis of the Ares I-X upper stage simulator (USS) element. As part of the design effort, the structural dynamic response of the Ares I-X launch vehicle to its vibroacoustic flight environments must be analyzed. The launch vehicle will be exposed to extremely high acoustic pressures during its lift-off and aerodynamic stages of flight. This in turn will cause high levels of random vibration on the vehicle's outer surface that will be transmitted to its interior. Critical flight equipment, such as its avionics and flight guidance components are susceptible to damage from this excitation. This study addresses the modelling, analysis and predictions from examining the structural dynamic response of the Ares I-X upper stage to its vibroacoustic excitations. A statistical energy analysis (SEA) model was used to predict the high frequency response of the vehicle at locations of interest. Key to this study was the definition of the excitation fields corresponding to lift off acoustics and the unsteady aerodynamic pressure fluctuations during flight. The predicted results will be used by the Ares I-X Project to verify the flight qualification status of the Ares I-X upper stage components.

Larko, Jeffrey M.↗
NASA NTRS2008

Induction Charge Detector with Multiple Sensing Stages

An induction charge detector with multiple sensing stages has been conceived for use in characterizing sprayed droplets, dust particles, large ionized molecules, and the like. Like related prior single-stage devices, each stage yields a measurement of the electric charge and the time of flight of the particle. In effect, an n-stage sensor yields n independent sets of such measurements from the same particle. The benefit of doing this is to increase the effective signal-to-noise ratio and thereby lower the charge-detection limit and the standard error of the charge measurement.

Gamero-Castano, Manuel↗
NASA NTRS2009

Advanced Development Program for a 625 lbf thrust engine for Ares First Stage Roll Control System

NASA's new Ares Launch Vehicle will require twelve thrusters to provide roll control of the vehicle during the first stage firing. All twelve roll control thrusters will be located at the inter-stage segment that separates the solid rocket booster first stage from the second stage. NASA selected a mono propellant hydrazine solution and as a result awarded Aerojet-General a contract in 2007 for an advanced development program for an MR-80- series 625 Ibf vacuum thrust monopropellant hydrazine thruster. This thruster has heritage dating back to the 1976 Viking Landers and most recently for the 2011 Mars Science Laboratory. Prior to the Ares application, the MR-80-series thrusters had been equipped with throttle valves and not typically operated in pulse mode. The primary objective of the advanced development program was to increase the technology readiness level and retire major technical risks for the future flight qualification test program. Aerojet built on their heritage MR-80 rocket engine designs to achieve the design and performance requirements. Significant improvements to cost and lead-time were achieved by applying Design for Manufacturing and Assembly (DFMA) principles. AerojetGeneral has completed Preliminary and Critical Design Reviews, followed by two successful rocket engine development test programs. The test programs included qualification random vibration and firing lite that significantly exceed the flight qualification requirements. This paper discusses the advanced development program and the demonstrated capability of the MR-80C engine. Y;

Dawson, Matt↗
NASA NTRS2010

NASA Ares I Launch Vehicle First Stage Roll Control System Cold Flow Development Test Program Overview

The Ares I launch vehicle is the selected design, chosen to return humans to the moon, Mars, and beyond. It is configured in two inline stages: the First Stage is a Space Shuttle derived five-segment Solid Rocket Booster and the Upper Stage is powered by a Saturn V derived J-2X engine. During launch, roll control for the First Stage (FS) is handled by a dedicated Roll Control System (RoCS) located on the connecting Interstage. That system will provide the Ares I with the ability to counteract induced roll torque while any induced yaw or pitch moments are handled by vectoring of the booster nozzle. This paper provides an overview of NASA s Ares I FS RoCS cold flow development test program including detailed test objectives, types of tests run to meet those objectives, an overview of the results, and applicable lessons learned. The test article was built and tested at the NASA Marshall Space Flight Center in Huntsville, AL. The FS RoCS System Development Test Article (SDTA) is a full scale, flight representative water flow test article whose primary objective was to obtain fluid system performance data to evaluate integrated system level performance characteristics and verify analytical models. Development testing and model correlation was deemed necessary as there is little historical precedent for similar large flow, pulsing systems such as the FS RoCS. The cold flow development test program consisted of flight-similar tanks, pressure regulators, and thruster valves, as well as plumbing simulating flight geometries, combined with other facility grade components and structure. Orifices downstream of the thruster valves were used to simulate the pressure drop through the thrusters. Additional primary objectives of this test program were to: evaluate system surge pressure (waterhammer) characteristics due to thruster valve operation over a range of mission duty cycles at various feed system pressures, evaluate temperature transients and heat transfer in the pressurization system, including regulator blowdown and propellant ullage performance, measure system pressure drops for comparison to analysis of tubing and components, and validate system activation and re-activation procedures for the helium pressurant system. Secondary objectives included: validating system processes for loading, unloading, and purging, validating procedures and system response for multiple failure scenarios, including relief valve operation, and evaluating system performance for contingency scenarios. The test results of the cold flow development test program are essential in validating the performance and interaction of the Roll Control System and anchoring analysis tools and results to a Critical Design Review level of fidelity.

Butt, Adam↗
NASA NTRS2010

Results of an Advanced Fan Stage Operating Over a Wide Range of Speed and Bypass Ratio: Comparison of CFD and Experimental Results - Part 2

NASA and GE teamed to design and build a 57 percent engine scaled fan stage for a Mach 4 variable cycle turbofan/ramjet engine for access to space with multipoint operations. This fan stage was tested in NASA's transonic compressor facility. The objectives of this test were to assess the aerodynamic and aero mechanic performance and operability characteristics of the fan stage over the entire range of engine operation including: 1) sea level static take-off; 2) transition over large swings in fan bypass ratio; 3) transition from turbofan to ramjet; and 4) fan wind-milling operation at high Mach flight conditions. This paper will focus on an assessment of APNASA, a multistage turbomachinery analysis code developed by NASA, to predict the fan stage performance and operability over a wide range of speeds (37 to 100 percent) and bypass ratios.

Celestina, Mark L.↗
NASA NTRS2011

First Stage Solid Propellant Multi Debris Thermal Analysis

The crew launch vehicle considered for the Constellation (Cx) Program utilizes a first stage solid rocket motor. If an abort is initiated in first stage flight the Crew Module (CM) will separate and be pulled away from the launch vehicle via a Launch Abort System (LAS) in order to safely and quickly carry the crew away from the malfunction launch vehicle. Having aborted the mission, the launch vehicle will likely be destroyed via a Flight Termination System (FTS) in order to prevent it from errantly traversing back over land and posing a risk to the public. The resulting launch vehicle debris field, composed primarily of first stage solid propellant, poses a threat to the CM. The harsh radiative thermal environment induced by surrounding burning propellant debris may lead to CM parachute failure. A methodology, detailed herein, has been developed to address this concern and quantify the risk of first stage propellant debris leading to radiative thermal demise of the CM parachutes. Utilizing basic thermal radiation principles, a software program was developed to calculate parachute temperature as a function of time for a given abort trajectory and debris piece trajectory set. Two test cases, considered worst-case aborts with regard to launch vehicle debris environments, were analyzed using the simulation: an abort declared at Mach 1 and an abort declared at maximum dynamic pressure (Max Q). For both cases, the resulting temperature profiles indicated that thermal limits for the parachutes were not exceeded. However, short duration close encounters by single debris pieces did have a significant effect on parachute temperature, with magnitudes on the order of 10 s of degrees Fahrenheit. Therefore while these two test cases did not indicate exceedance of thermal limits, in order to quantify the risk of parachute failure due to radiative effects from the abort environment, a more thorough probability-based analysis using the methodology demonstrated herein must be performed.

Toleman, Benjamin M.↗
NASA NTRS2011

Cryo Propulsive Stage: HEFT Phase 2 Point of Departure

The CPS is an in-space high thrust propulsive stage based largely on state of the practice design for launch vehicle upper stages. However, unlike conventional propulsive stages, it also contains power generation and thermal control systems to limit the loss of liquid hydrogen and oxygen due to boil-off during extended in-space storage. The CPS provides high thrust GV for rapid transfer of in-space elements to their destinations or staging points (i.e., E-M L1). The CPS is designed around a block upgrade strategy to provide maximum mission/architecture flexibility: a) Block 1 CPS: Short duration flight times (hours), passive cryofluid management. b) Block 2 CPS: Long duration flight times (days/weeks/months), active and passive cryofluid management.

Martinez, Roland M.↗
NASA NTRS2011

First Stage Solid Propellant Multiply Debris Thermal Analysis

Destruction of a solid rocket stage of a launch vehicle can create a thermal radiation hazard for an aborting crew module. This hazard was assessed for the Constellation Program (Cx) crew and launch vehicle concept. For this concept, if an abort was initiated in first stage flight, the Crew Module (CM) will separate and be pulled away from the malfunctioning launch vehicle via a Launch Abort System (LAS). Having aborted the mission, the launch vehicle will likely be destroyed via a Flight Termination System (FTS) in order to prevent it from errantly traversing back over land and posing a risk to the public. The resulting launch vehicle debris field, composed primarily of first stage solid propellant, poses a threat to the CM. The harsh radiative thermal environment, caused by surrounding burning propellant debris, may lead to CM parachute failure. A methodology, detailed herein, has been developed to address this concern and to quantify the risk of first stage propellant debris leading to the thermal demise of the CM parachutes. Utilizing basic thermal radiation principles, a software program was developed to calculate parachute temperature as a function of time for a given abort trajectory and debris piece trajectory set. Two test cases, considered worst case aborts with regard to launch vehicle debris environments, were analyzed using the simulation: an abort declared at Mach 1 and an abort declared at maximum dynamic pressure (Max Q). For both cases, the resulting temperature profiles indicated that thermal limits for the parachutes were not exceeded. However, short duration close encounters by single debris pieces did have a significant effect on parachute temperature. Therefore while these two test cases did not indicate exceedance of thermal limits, in order to quantify the risk of parachute failure due to radiative effects from the abort environment, a more thorough probability-based analysis using the methodology demonstrated herein must be performed.

Toleman, Benjamin M.↗
NASA NTRS2010

Stage Separation Failure: Model Based Diagnostics and Prognostics

Safety of the next-generation space flight vehicles requires development of an in-flight Failure Detection and Prognostic (FD&P) system. Development of such system is challenging task that involves analysis of many hard hitting engineering problems across the board. In this paper we report progress in the development of FD&P for the re-contact fault between upper stage nozzle and the inter-stage caused by the first stage and upper stage separation failure. A high-fidelity models and analytical estimations are applied to analyze the following sequence of events: (i) structural dynamics of the nozzle extension during the impact; (ii) structural stability of the deformed nozzle in the presence of the pressure and temperature loads induced by the hot gas flow during engine start up; and (iii) the fault induced thrust changes in the steady burning regime. The diagnostic is based on the measurements of the impact torque. The prognostic is based on the analysis of the correlation between the actuator signal and fault-induced changes in the nozzle structural stability and thrust.

Luchinsky, Dmitry↗
NASA NTRS2011

Cryogenic Propulsion Stage

The CPS is an in-space cryogenic propulsive stage based largely on state of the practice design for launch vehicle upper stages. However, unlike conventional propulsive stages, it also contains power generation and thermal control systems to limit the loss of liquid hydrogen and oxygen due to boil-off during extended in-space storage. The CPS provides the necessary (Delta)V for rapid transfer of in-space elements to their destinations or staging points (i.e., E-M L1). The CPS is designed around a block upgrade strategy to provide maximum mission/architecture flexibility. Block 1 CPS: Short duration flight times (hours), passive cryo fluid management. Block 2 CPS: Long duration flight times (days/weeks/months), active and passive cryo fluid management.

Jones, David↗
NASA NTRS2011

ARES I Upper Stage Subsystems Design and Development

From 2005 through early 2011, NASA conducted concept definition, design, and development of the Ares I launch vehicle. The Ares I was conceived to serve as a crew launch vehicle for beyond-low-Earth-orbit human space exploration missions as part of the Constellation Program Architecture. The vehicle was configured with a single shuttle-derived solid rocket booster first stage and a new liquid oxygen/liquid hydrogen upper stage, propelled by a single, newly developed J-2X engine. The Orion Crew Exploration Vehicle was to be mated to the forward end of the Ares I upper stage through an interface with fairings and a payload adapter. The vehicle design passed a Preliminary Design Review in August 2008, and was nearing the Critical Design Review when efforts were concluded as a result of the Constellation Program s cancellation. At NASA Glenn Research Center, four subsystems were developed for the Ares I upper stage. These were thrust vector control (TVC) for the J-2X, electrical power system (EPS), purge and hazardous gas (P&HG), and development flight instrumentation (DFI). The teams working each of these subsystems achieved 80 percent or greater design completion and extensive development testing. These efforts were extremely successful representing state-of-the-art technology and hardware advances necessary to achieve Ares I reliability, safety, availability, and performance requirements. This paper documents the designs, development test activity, and results.

Frate, David T.↗
NASA NTRS2011

Methane and Environmental Change during the Paleocene-Eocene Thermal Maximum (PETM): Modeling the PETM Onset as a Two-stage Event

An atmospheric CH4 box model coupled to a global carbon cycle box model is used to constrain the carbon emission associated with the PETM and assess the role of CH4 during this event. A range of atmospheric and oceanic emission scenarios representing different amounts, rates, and isotopic signatures of emitted carbon are used to model the PETM onset. The first 3 kyr of the onset, a pre-isotope excursion stage, is simulated by the atmospheric release of 900 to 1100 Pg C CH4 with a delta C-13 of -22 to - 30 %. For a global average warming of 3 deg C, a release of CO2 to the ocean and CH4 to the atmosphere totalling 900 to 1400 Pg C, with a delta C-13 of -50 to -60%, simulates the subsequent 1 -kyr isotope excursion stage. To explain the observations, the carbon must have been released over at most 500 years. The first stage results cannot be associated with any known PETM hypothesis. However, the second stage results are consistent with a methane hydrate source. More than a single source of carbon is required to explain the PETM onset.

Carozza, David A.↗
NASA NTRS2012

End-To-End Simulation of Launch Vehicle Trajectories Including Stage Separation Dynamics

The development of methodologies, techniques, and tools for analysis and simulation of stage separation dynamics is critically needed for successful design and operation of multistage reusable launch vehicles. As a part of this activity, the Constraint Force Equation (CFE) methodology was developed and implemented in the Program to Optimize Simulated Trajectories II (POST2). The objective of this paper is to demonstrate the capability of POST2/CFE to simulate a complete end-to-end mission. The vehicle configuration selected was the Two-Stage-To-Orbit (TSTO) Langley Glide Back Booster (LGBB) bimese configuration, an in-house concept consisting of a reusable booster and an orbiter having identical outer mold lines. The proximity and isolated aerodynamic databases used for the simulation were assembled using wind-tunnel test data for this vehicle. POST2/CFE simulation results are presented for the entire mission, from lift-off, through stage separation, orbiter ascent to orbit, and booster glide back to the launch site. Additionally, POST2/CFE stage separation simulation results are compared with results from industry standard commercial software used for solving dynamics problems involving multiple bodies connected by joints.

Albertson, Cindy W.↗
NASA NTRS1944

Altitude Performance of Allison V-1710-93 Engine and Two-Stage Supercharger

Bench tests have been made of an Allison V-1710-93 engine at simulated altitudes of 17,000 and 29,000 feet to determine the performance of the engine, the characteristics of the induction system, and the modifications required in the induction system for the center dot development of 1840 brake horsepower at 17,030 feet and 1200 horsepower at 29,000 feet. With the engine speed maintained at a constant value of 3000 rpm, the tip speed of the auxiliary-stage supercharger was varied from the minimum to the maximum speed. Measurements were made to determine engine power, fuel and air consumptions, and pressure and temperature changes introduced by each component of the induction system. Although the over-all pressure ratios of the superchargers agreed with those determined from previous sea-level bench tests of these units, the over-all temperature rise was -generally much lower than had been anticipated. In order to obtain the desired brake horsepower, some form of charge-air cooling is essential; the carburetor should either be enlarged or removed; and the minimum gear ratios of the engine-stage and the minimum gear ratios of the engine-stage and the auxiliary-stage superchargers should be 8.1:1 and 8.08:1, respectively.

Edward Glodeck↗
NASA NTRS2015

Thermodynamic Analysis of the 3-Stage ADR for the Astro-H Soft X-Ray Spectrometer Instrument

The Soft X-ray Spectrometer (SXS) instrument on Astro-H will use a 3-stage ADR to cool the microcalorimeter array to 50 mK. In the primary operating mode, two stages of the ADR cool the detectors using superfluid helium at 1.20 K as the heat sink. In the secondary mode, which is activated when the liquid helium is depleted, two of the stages continuously cool the (empty) helium tank using a 4.5 K Joule-Thomson cooler as the heat sink, and the third stage cools the detectors. In the design phase, a high-fidelity model of the ADR was developed in order to predict both the cooling capacity and heat rejection rates in both operating modes. The primary sources of heat flow are from the salt pills, hysteresis heat from the magnets and magnetic shields, and power dissipated by the heat switches. The flight instrument dewar, ADR, detectors and electronics were integrated in mid-2014 and have since undergone extensive performance testing, in part to validate the performance model. This paper will present the thermodynamic performance of the ADR, including cooling capacity, heat rejection to the heat sinks, and various measures of efficiency.

space cryogenics↗