Search NASA⌕ Search

SEARCH · Search NASA

Results for “ENGINES, ROCKET - ALTITUDE TESTS”

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.

Quote a phrase for an exact phrase match. Source license links do not imply unrestricted reuse.

At least 181 records · Page 10
All sourcesNASA NTRSDOE OSTI
NASA NTRS1991

The development of a post-test diagnostic system for rocket engines

An effort was undertaken by NASA to develop an automated post-test, post-flight diagnostic system for rocket engines. The automated system is designed to be generic and to automate the rocket engine data review process. A modular, distributed architecture with a generic software core was chosen to meet the design requirements. The diagnostic system is initially being applied to the Space Shuttle Main Engine data review process. The system modules currently under development are the session/message manager, and portions of the applications section, the component analysis section, and the intelligent knowledge server. An overview is presented of a rocket engine data review process, the design requirements and guidelines, the architecture and modules, and the projected benefits of the automated diagnostic system.

Zakrajsek, June F.↗
NASA NTRS2011

Summary of Results from Space Shuttle Main Engine Off-Nominal Testing

This paper is a summary of Space Shuttle Main Engine (SSME) off-nominal testing that occurred during 2008 and 2009. During the last two years of planned SSME testing at Stennis Space Center, Pratt & Whitney Rocketdyne worked with their NASA MSFC customer to systematically identify, develop, assess, and implement challenging test objectives in order to expand the knowledge of one of the world s most reliable and highly tested large rocket engine. The objectives successfully investigated three main areas of interest expanding engine performance margins, demonstrating system operational capabilities, and establishing ground work for new rocket engine technology. The testing gave the Space Shuttle Program new options to safely fly out the flight manifest and provided Pratt & Whitney Rocketdyne and NASA new insight into the operational capabilities of the SSME, capabilities which can be used in assessing potential future applications of the RS-25 engine.

Horton, James F.↗
NASA NTRS1991

Rover nuclear rocket engine program: Overview of rover engine tests

The results of nuclear rocket development activities from the inception of the ROVER program in 1955 through the termination of activities on January 5, 1973 are summarized. This report discusses the nuclear reactor test configurations (non cold flow) along with the nuclear furnace demonstrated during this time frame. Included in the report are brief descriptions of the propulsion systems, test objectives, accomplishments, technical issues, and relevant test results for the various reactor tests. Additionally, this document is specifically aimed at reporting performance data and their relationship to fuel element development with little or no emphasis on other (important) items.

Finseth, J. L.↗
NASA NTRS2018

Initial Criticality Assessments to Guide FMEAs on Rocket Engine Hot Fire Testing

Presented is an explanation of the use of the Initial Criticality Assessment (ICA) technique, a triage process for prioritizing required Failure Modes and Effects Analysis (FMEAs), for the European Service Module's main propulsion system's hot-fire test bed at White Sands New Mexico. Rather than instinctively performing many FMEAs of subsystems, or one large system level FMEA where every subcomponent is analyzed, the ICA guided an informed analysis of only the hardware that had a large impact to hazards. The low criticality hardware was documented via the ICA and no FMEA was performed; the work could then focus on the high criticality hardware. Thus a savings of Program resources was achieved. The experiences gained in creating these ICAs for this international collaborative project confirmed that the need for continuous communication across the technical teams is one of the greatest areas of emphasis. The European Space Agency (ESA) is developing the European Service Module (ESM), with its primary contractor, Airbus Defence and Space in Germany, for delivery to the National Aeronautics and Space Administration (NASA). The module will be equipped with a total of 21 engines to support NASA’s Orion spacecraft: one U.S. Space Shuttle Orbital Maneuvering System-Engine (OMS-E), eight auxiliary thrusters and 12 smaller RCS (Reaction Control System) thrusters. The main ESM propulsion system, used for large translational maneuvers, consists of one OMS-E... To qualify the design of the ESM propulsion subsystem (PSS) an all-steel Propulsion Qualification Module (PQM) structure is used to test the propulsion systems on Orion, including “hot firing” of the OMS engine, thrusters, and RCS. The PQM has been developed as a hot-fire test bed to be tested at the NASA White Sands Test Facility (WSTF). One of the objectives of the testing is to assure that the OMS-E can be safely operated with the PQM. Testing will also demonstrate that the PQM can set the proper upstream pressures and temperatures for the OMS-E to operate nominally given the PQM has never been tested in hot-fire operation with OMS-E before. In order to safely conduct the test campaign, hardware such as the engine subassembly, fluid feed lines, valves, electrical power lines, instrumentation, stiff links, installation Ground Support Equipment (GSE), and diffuser [whose objectives are to collect the exhaust of the OMS-E to actively cool down the exhaust gases, reduce thermal exchanges, and create a vacuum at OMS-E level before igniting], had to be analyzed for any hazards and failure modes.

Fault Tolerance↗
NASA NTRS2004

Multi-Element Unstructured Analyses of Complex Valve Systems

The safe and reliable operation of high pressure test stands for rocket engine and component testing places an increased emphasis on the performance of control valves and flow metering devices. In this paper, we will present a series of high fidelity computational analyses of systems ranging from cryogenic control valves and pressure regulator systems to cavitating venturis that are used to support rocket engine and component testing at NASA Stennis Space Center. A generalized multi-element framework with sub-models for grid adaption, grid movement and multi-phase flow dynamics has been used to carry out the simulations. Such a framework provides the flexibility of resolving the structural and functional complexities that are typically associated with valve-based high pressure feed systems and have been difficult to deal with traditional CFD methods. Our simulations revealed a rich variety of flow phenomena such as secondary flow patterns, hydrodynamic instabilities, fluctuating vapor pockets etc. In the paper, we will discuss performance losses related to cryogenic control valves, and provide insight into the physics of the dominant multi-phase fluid transport phenomena that are responsible for the choking like behavior in cryogenic control elements. Additionally, we will provide detailed analyses of the modal instability that is observed in the operation of the dome pressure regulator valve. Such instabilities are usually not localized and manifest themselves as a system wide phenomena leading to an undesirable chatter at high flow conditions.

Sulyma, Peter↗
NASA NTRS2004

A Method for Calculating the Probability of Successfully Completing a Rocket Propulsion Ground Test

Propulsion ground test facilities face the daily challenges of scheduling multiple customers into limited facility space and successfully completing their propulsion test projects. Due to budgetary and schedule constraints, NASA and industry customers are pushing to test more components, for less money, in a shorter period of time. As these new rocket engine component test programs are undertaken, the lack of technology maturity in the test articles, combined with pushing the test facilities capabilities to their limits, tends to lead to an increase in facility breakdowns and unsuccessful tests. Over the last five years Stennis Space Center's propulsion test facilities have performed hundreds of tests, collected thousands of seconds of test data, and broken numerous test facility and test article parts. While various initiatives have been implemented to provide better propulsion test techniques and improve the quality, reliability, and maintainability of goods and parts used in the propulsion test facilities, unexpected failures during testing still occur quite regularly due to the harsh environment in which the propulsion test facilities operate. Previous attempts at modeling the lifecycle of a propulsion component test project have met with little success. Each of the attempts suffered form incomplete or inconsistent data on which to base the models. By focusing on the actual test phase of the tests project rather than the formulation, design or construction phases of the test project, the quality and quantity of available data increases dramatically. A logistic regression model has been developed form the data collected over the last five years, allowing the probability of successfully completing a rocket propulsion component test to be calculated. A logistic regression model is a mathematical modeling approach that can be used to describe the relationship of several independent predictor variables X(sub 1), X(sub 2),..,X(sub k) to a binary or dichotomous dependent variable Y, where Y can only be one of two possible outcomes, in this case Success or Failure. Logistic regression has primarily been used in the fields of epidemiology and biomedical research, but lends itself to many other applications. As indicated the use of logistic regression is not new, however, modeling propulsion ground test facilities using logistic regression is both a new and unique application of the statistical technique. Results from the models provide project managers with insight and confidence into the affectivity of rocket engine component ground test projects. The initial success in modeling rocket propulsion ground test projects clears the way for more complex models to be developed in this area.

Messer, Bradley P.↗
NASA NTRS1974

A study of the durability of beryllium rocket engines

An experimental test program was performed to demonstrate the durability of a beryllium INTEREGEN rocket engine when operating under conditions simulating the space shuttle reaction control system. A vibration simulator was exposed to the equivalent of 100 missions of X, Y, and Z axes random vibration to demonstrate the integrity of the recently developed injector-to-chamber braze joint. An off-limits engine was hot fired under extreme conditions of mixture ratio, chamber pressure, and orifice plugging. A durability engine was exposed to six environmental cycles interspersed with hot-fire tests without intermediate cleaning, service, or maintenance. Results from this program indicate the ability of the beryllium INTEREGEN engine concept to meet the operational requirements of the space shuttle reaction control system.

Paster, R. D.↗
NASA NTRS2003

Nonlinear Modeling and Control of a Propellant Mixer

A mixing chamber used in rocket engine combustion testing at NASA Stennis Space Center is modeled by a second order nonlinear MIMO system. The mixer is used to condition the thermodynamic properties of cryogenic liquid propellant by controlled injection of the same substance in the gaseous phase. The three inputs of the mixer are the positions of the valves regulating the liquid and gas flows at the inlets, and the position of the exit valve regulating the flow of conditioned propellant. The outputs to be tracked and/or regulated are mixer internal pressure, exit mass flow, and exit temperature. The outputs must conform to test specifications dictated by the type of rocket engine or component being tested downstream of the mixer. Feedback linearization is used to achieve tracking and regulation of the outputs. It is shown that the system is minimum-phase provided certain conditions on the parameters are satisfied. The conditions are shown to have physical interpretation.

Barbieri, Enrique↗
NASA NTRS2016

Subscale Carbon-Carbon Nozzle Extension Development and Hot Fire Testing in Support of Upper Stage Liquid Rocket Engines

Upper stage and in-space liquid rocket engines are optimized for performance through the use of high area ratio nozzles to fully expand combustion gases to low exit pressures increasing exhaust velocities. Due to the large size of such nozzles and the related engine performance requirements, carbon-carbon (C/C) composite nozzle extensions are being considered for use in order to reduce weight impacts. NASA and industry partner Carbon-Carbon Advanced Technologies (C-CAT) are working towards advancing the technology readiness level of large-scale, domestically-fabricated, C/C nozzle extensions. These C/C extensions have the ability to reduce the overall costs of extensions relative to heritage metallic and composite extensions and to decrease weight by 50%. Material process and coating developments have advanced over the last several years, but hot fire testing to fully evaluate C/C nozzle extensions in relevant environments has been very limited. NASA and C-CAT have designed, fabricated and hot fire tested multiple subscale nozzle extension test articles of various C/C material systems, with the goal of assessing and advancing the manufacturability of these domestically producible materials as well as characterizing their performance when subjected to the typical environments found in a variety of liquid rocket and scramjet engines. Testing at the MSFC Test Stand 115 evaluated heritage and state-of-the-art C/C materials and coatings, demonstrating the capabilities of the high temperature materials and their fabrication methods. This paper discusses the design and fabrication of the 1.2k-lbf sized carbon-carbon nozzle extensions, provides an overview of the test campaign, presents results of the hot fire testing, and discusses potential follow-on development work.

Gradl, Paul↗
NASA NTRSDate not supplied

Large Vehicle Lunar Landing Surface Interaction and In-Situ Resource Based Risk Mitigation: Landing & Launch Pads

A key capability required for the exploration of planetary bodies is the ability to land on the surface. Previous work performed by NASA and other institutions has primarily focused on landing small spacecraft on planetary surfaces and the associated small-to-medium thrusters required for the soft landing. In the case of human exploration—particularly the establishment of long duration exploration and habitation outposts—the ability to land large landers, such as the SpaceX Starship, is necessary. These larger landing systems require the use of more powerful engines, with higher engine exhaust temperatures and higher landing loads. Understanding the excavation of material by the engines, as well as the potential for the landing legs to sink into the subsurface, is key in ensuring reliable and safe landings. A further improvement in landing reliability can be achieved by constructing landing / launch pads, especially with in-situ resources. Some material excavation by the plume is inevitable, leaving at least a portion of the surface scoured and uneven under the lander and ejecting regolith particles and rocks at very high velocities. One possible solution would be to robotically build landing / launch pads (ideally autonomously) at the destination using in-situ materials. In this case, the first one or few landers will need to land on unimproved surfaces at higher risk; however, they would bring the required equipment to build the landing pads with mostly local resources, thus increasing the reliability of safe landing for subsequent larger landers. A number of methods to build in-situ landing and launch pads have already been developed. These methods include no or some addition of required binder additives to the local regolith material, different processing approaches and result in varying landing pad strengths. A sub-scale rocket engine plume, was used to simulate some of the conditions of a landing on the Moon to assess the effectiveness of various materials for an in-situ built landing pad, The GO2/GCH4 rocket engine fired on 1m2 area coupons of representative pad materials. The results will allow continued development towards materials that satisfy the landing pad properties required for the effective risk reduction and increased reliability for landing people and equipment on the lunar surface. This work contained two parts: (1) computer modeling of a large rocket engine plume interacting with regolith on the Moon, using the Granular Gas Flow Solver (GGFS) provided by CFD Research Corporation as well as other computational fluid dynamics codes (CFD) such as Loci/CHEM. (2) Developing landing/launch pad materials that could be used for in-situ construction on the lunar surface in the future, to mitigate the calculated effects of a large vehicle rocket engine landing and launching on the Moon. The resulting computed values of plume impingement surface temperature, stagnation pressure, gas velocity, shear stress and heat flux were then matched as closely as possible in the Earth’s atmosphere in a sub-scale rocket engine GO2/CH4 test which was provided by Masten Space Systems in Mojave, California. The rocket engine was mounted on a test stand with vertical translation capabilities so that the landing operations of a lander could be simulated (Figure 1).The pad materials test coupons were placed at the surface of a large bin containing simulated lunar basalt regolith and subjected to a test firing as shown in Figure 2. The results of this testing will be presented with related findings, analysis and discussion.

Plume Surface Interaction↗
NASA NTRS2007

Fiber Optic Raman Sensor to Monitor Concentration Ratio of Nitrogen and Oxygen in a Cryogenic Mixture

A spontaneous Raman scattering optical fiber sensor is developed for a specific need of NASA/SSC for long-term detection and monitoring of the quality of liquid oxygen (LOX) in the delivery line during ground testing of rocket engines. The sensor performance was tested in the laboratory and with different excitation light sources. To evaluate the sensor performance with different excitation light sources for the LOX quality application, we have used the various mixtures of liquid oxygen and liquid nitrogen as samples. The study of the sensor performance shows that this sensor offers a great deal of flexibility and provides a cost effective solution for the application. However, an improved system response time is needed for the real-time, quantitative monitoring of the quality of cryogenic fluids in harsh environment.

Tiwari, Vidhu S.↗
NASA NTRS1992

Multiple dopant injection system for small rocket engines

The Diagnostics Test Facility (DTF) at NASA's Stennis Space Center (SSC) was designed and built to provide a standard rocket engine exhaust plume for use in the research and development of engine health monitoring instrumentation. A 1000 lb thrust class liquid oxygen (LOX)-gaseous hydrogen (GH2) fueled rocket engine is used as the subscale plume source to simulate the SSME during experimentation and instrument development. The ability of the DTF to provide efficient, and low cost test operations makes it uniquely suited for plume diagnostic experimentation. The most unique feature of the DTF is the Multiple Dopant Injection System (MDIS) that is used to seed the exhaust plume with the desired element or metal alloy. The dopant injection takes place at the fuel injector, yielding a very uniform and homogeneous distribution of the seeding material in the exhaust plume. The MDIS allows during a single test firing of the DTF, the seeding of the exhaust plume with up to three different dopants and also provides distilled water base lines between the dopants. A number of plume diagnostic-related experiments have already utilized the unique capabilities of the DTF.

Sakala, G. G.↗
NASA NTRSDate not supplied

Development and Testing of Pulse Guns for Combustion Instability Testing

Combustion stability requirements of a liquid propellant rocket engine are most commonly verified by evaluating the oscillatory pressure or vibration characteristics following an artificial disturbance of the combustion in the combustion chamber. Two types of devices are most commonly used for dynamic stability rating – nondirectional bombs and pulse guns. Bombs have historically been used most often, especially for the large engines developed around the Apollo era. However, bombs in recent years have become much more expensive and difficult to procure, and their transportation and handling requirements have become more demanding. Pulse guns, with inherently less expensive materials, and far less restrictive handling and shipping requirements, have become an attractive alternative. The NASA Marshall Space Flight Center has developed a pulse gun for use during future development programs at the center, and as a design reference for interested commercial users. The objective of this paper is to present the different variants of the pulse gun and characterize performance over a range of parameters. The parameters include the breech diameter and corresponding volume, barrel length, amount of gun powder used, the method of containing the gunpowder within the breech, and the pressure rating of the burst disk. In total, 44 tests were conducted with the pulse gun firing into a test chamber pressurized to 2300 psig with gaseous nitrogen. The best experimental results – including the highest and most consistent peak pressure amplitudes recorded by dynamic pressure sensors in the test chamber – were from a configuration consisting of a 0.40” inner-diameter breech, 15-16 grains of gunpowder wrapped in cigarette paper, and a 24,000 psid burst disk. With this configuration, the zero-to-peak overpressures recorded in the test chamber ranged from about 37% to 58% of the 2300 psig mean pressure, which is adequate for typical combustion stability rating. To complement the data analysis, the acoustics of the pressure waves as they propagated through the pulse gun and into the test chamber were modeled. A more comprehensive final report of this pulse gun development program is planned to be available in the near future.

Pulse Gun↗
NASA NTRS2009

Hydrogen Reclamation and Reutilization

John C. Stennis Space Center (SSC) provides rocket engine propulsion testing for NASA's space programs. Since the development of the Space Shuttle, every Space Shuttle Main Engine (SSME) has undergone acceptance testing at SSC before going to Kennedy Space Center (KSC) for integration into the Space Shuttle. The SSME is a large cryogenic rocket engine that uses Liquid Hydrogen (LH2) as the fuel. As NASA moves to the new ARES V launch system, the main engines on the new vehicle, as well as the upper stage engine, are currently base lined to be cryogenic rocket engines that will also use LH2. The main rocket engines for the ARES V will be larger than the SSME, while the upper stage engine will be approximately half that size. As a result, significant quantities of hydrogen will be required during the development, testing, and operation of these rocket engines.

Hebert, Bartt↗