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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 73 records · Page 4
All sourcesNASA NTRSDOE OSTI
NASA NTRS1985

Rotor instability due to loose rotating part

Loosening of a rotating part from its fixed position on the shaft or a part of the stator which comes loose and begins to turn with the rotor very frequently represents machinery malfunction. The loose part becomes involved in rotative motion mostly due to dry or fluid friction, and thus its motion is very erratic. The loose part can also move axially along the shaft. Detachment of the rotating part causes changes in the rotor balance state. Most often this results in higher unbalance. During steady-state operation the effect of a loose rotating part can manifest itself through heat vibration. It can be diagnosed by observing periodic changes of amplitude and phase of the synchronous response. During start-up (or shutdown) a loose rotating part carrying some amount of unbalance may manifest its dynamic action in the form of subsynchronous vibrations, very similar to those of other instabilities. The objective of this demonstration is to observe the effect of a loose rotating part (fixed, however, in the axial direction) under both steady-state (rotor constant speed) and transient (rotor start-up or shutdown) operation. The dynamic response depends very much on the amount of damping in the system: lubrication of the loose part/shaft surfaces and addition/elimination of aerodynamic drag blades, mounted on the loose disk, significantly change the rotor response.

Muszynska, A.↗
NASA NTRS1987

Probabilistic risk analysis of flying the space shuttle with and without fuel turbine discharge temperature redline protection

An exact mathematical model is presented of the Space Shuttle Main Engine computer voting logic in the presence of High Pressure Fuel Turbine (HPFT) overtemperature events and fuel turbine temperature sensor failures. The model provides estimates of the probability of erroneous engine shutdown and the probability of not detecting a HPFT overtemperature event. Because it is believed that the likelihood of sensor failures and overtemperature events in the HPFT greatly overshadows those in the High Pressure Oxygen Turbine (HPOT), this modeling effort focused on the HPFT. However, because the redline protection logic is the same for both turbines, estimation of the model parameters using relevant HPOT data would provide estimates of the risk of erroneous engine shutdown and an undetected overtemperature event in the HPOT. Because of the complexity of the model, it was necessary to program the solution which thus makes it feasible to accommodate a changing data base. This is considered to be of great interest because of the subjective nature in determining the relevancy of certain failures and the fact that the data base is constantly changing as a result of the frequent engine tests.

Howell, Leonard↗
NASA NTRS1987

Computer controlled operation of a two-engine xenon ion propulsion system

The development and testing of a computer control system for a two-engine xenon ion propulsion module is described. The computer system controls all aspects of the propulsion module operation including: start-up, steady-state operation, throttling and shutdown of the engines; start-up, operation and shutdown of the central neutralizer subsystem; control of the gimbal system for each engine; and operation of the valves in the propellant storage and distribution system. The most important engine control algorithms are described in detail. These control algorithms provide flexibility in the operation and throttling of ion engines which has never before been possible. This flexibility is made possible in large part through the use of flow controllers which maintain the total flow rate of propellant into the engine at the proper level. Data demonstrating the throttle capabilities of the engine and control system are presented.

Brophy, John R.↗
NASA NTRS1986

Transient modeling of the thermohydraulic behavior of high temperature heat pipes for space reactor applications

Many proposed space reactor designs employ heat pipes as a means of conveying heat. Previous researchers have been concerned with steady state operation, but the transient operation is of interest in space reactor applications due to the necessity of remote startup and shutdown. A model is being developed to study the dynamic behavior of high temperature heat pipes during startup, shutdown and normal operation under space environments. Model development and preliminary results for a hypothetical design of the system are presented.

Hall, Michael L.↗
NASA NTRS1987

Pattern recognition techniques for failure trend detection in SSME ground tests

The Space Shuttle Main Engine (SSME) is a complex power plant. To evaluate its performance 1200 hot-wire ground tests have been conducted, varying in duration from 0 to 500 secs. During the test some 500 sensors are sampled every 20 ms. The sensors are generally bounded by red lines so that an excursion beyond could lead to premature shutdown. In 27 tests it was not possible to effect an orderly premature shutdown, resulting in major incidents with serious damage to the SSME and test stand. The application of pattern recognition are investigated to detect SSME performance trends that may lead to major incidents. Based on the sensor data a set of (n) features is defined. At any time during the test, the state of the SSME is given by a point in the n-dimensional feature space. The history of a test can now be represented as a trajectory in the n-dimensional feature space. Portions of the normal trajectories and failed test trajectories would lie in different regions of the n-dimensional feature space. The latter can now be partitioned into regions of normal and failed tests. Thus, it is possible to examine the trajectory of a test in progress and predict if it is going into the normal or failure region.

Choudry, A.↗
NASA NTRS1986

Evaluation and Improvement of Liquid Propellant Rocket Chugging Analysis Techniques. Part 1: A One-Dimensional Analysis of Low Frequency Combustion Instability in the Fuel Preburner of the Space Shuttle Main Engine

Low frequency combustion instability, known as chugging, is consistently experienced during shutdown in the fuel and oxidizer preburners of the Space Shuttle Main Engines. Such problems always occur during the helium purge of the residual oxidizer from the preburner manifolds during the shutdown sequence. Possible causes and triggering mechanisms are analyzed and details in modeling the fuel preburner chug are presented. A linearized chugging model, based on the foundation of previous models, capable of predicting the chug occurrence is discussed and the predicted results are presented and compared to experimental work performed by NASA. Sensitivity parameters such as chamber pressure, fuel and oxidizer temperatures, and the effective bulk modulus of the liquid oxidizer are considered in analyzing the fuel preburner chug. The computer program CHUGTEST is utilized to generate the stability boundary for each sensitivity study and the region for stable operation is identified.

Lim, Kair Chuan↗
NASA NTRS1988

Development of an advanced failure detection algorithm for the SSME

Since ground testing of the Space Shuttle Main Engine began in 1975, a multifaceted detection/shutdown system has been used to detect engine anomalies and initiate shutdown. In spite of this system, 27 major incidents have occurred since testing began. Although a small percentage of over 1200 successful hot fire tests, such incidents can result in significant cost and schedule impacts. Therefore, an advanced failure detection algorithm has been developed which consists of three detection approaches designed for a failure characteristic that has been previously documented. Simulations using sensor data from five incident tests show the algorithm detecting an anomally and signalling shutdown prior to the redline cutoff time. The foundation elements of the algorithm and its functioning are described.

Norman, A.↗
NASA NTRS1989

Effect of Correlation on Multi-Engine Rocket Propulsion Systems

A matter of great concern in the design and operation of multi-engine rocket propulsion systems is the effect of the premature shutdown of one engine on the vehicle. This probability that a premature shutdown will cause a vehicle loss is termed correlation. Based on airbreathing experiences as well as rocket engine data the best estimate of this correlation is made and then applied to the overall multi-engine reliability problem to demonstrate its potential effect. At this point, follow-on analyses are pointed out that illustrate how any potential failures that may cause a correlatable event can be eliminated; thus bringing this correlation to almost 0.

Baker, J. R.↗
NASA NTRS1990

Transportation monitoring unit qualification

Transportation monitoring unit (TMU) qualification testing was performed between 3 Mar. and 14 Dec. 1989. The purpose of the testing was to qualify the TMUs to monitor and store temperature and acceleration data on redesigned solid rocket motor segments and exit cones while they are being shipped from Utah's Thiokol Corporation, Space Operations, to Kennedy Space Center. TMUs were subjected to transportation tests that concerned the structural integrity of the TMUs only, and did not involve TMU measuring capability. This testing was terminated prior to completion due to mounting plate failures, high and low temperature shutdown failures, and data collection errors. Corrective actions taken by the vendor to eliminate high temperature shutdowns were ineffective. An evaluation was performed on the TMUs to determine the TMU vibration and temperature measuring accuracy at a variety of temperatures. This test demonstrated that TMU measured shock levels are high, and that TMUs are temperature sensitive because of decreased accuracy at high and low temperatures. It was determined that modifications to the current TMU system, such that it could be qualified for use, would require a complete redesign and remanufacture. Because the cost of redesigning and remanufacturing the present TMU system exceeds the cost of procuring a new system that could be qualified without modification, it is recommended that an alternate transportation monitoring system be qualified.

Cook, M.↗
NASA NTRS1990

Health Monitoring System for the SSME-fault detection algorithms

A Health Monitoring System (HMS) Framework for the Space Shuttle Main Engine (SSME) has been developed by United Technologies Corporation (UTC) for the NASA Lewis Research Center. As part of this effort, fault detection algorithms have been developed to detect the SSME faults with sufficient time to shutdown the engine. These algorithms have been designed to provide monitoring coverage during the startup, mainstage and shutdown phases of the SSME operation. The algorithms have the capability to detect multiple SSME faults, and are based on time series, regression and clustering techniques. This paper presents a discussion of candidate algorithms suitable for fault detection followed by a description of the algorithms selected for implementation in the HMS and the results of testing these algorithms with the SSME test stand data.

Tulpule, S.↗
NASA NTRS1990

Douglas flight deck design philosophy

The systems experience gained from 17 years of DC-10 operation was used during the design of the MD-11 to automate system operation and reduce crew workload. All functions, from preflight to shutdown at the termination of flight, require little input from the crew. The MD-11 aircraft systems are monitored for proper operation by the Aircraft Systems Controllers (ASC). In most cases, system reconfiguration as a result of a malfunction is automated. Manual input is required for irreversible actions such as engine shutdown, fuel dump, fire agent discharge, or Integrated Drive Generator (IDG) disconnect. During normal operations, when the cockpit is configured for flight, all annunciators on the overhead panel will be extinguished. This Dark Cockpit immediately confirms to the crew that the panels are correctly configured and that no abnormalities are present. Primary systems annunciations are shown in text on the Alert Area of the Engine and Alert Display (EAD). This eliminates the need to scan the overhead. The MD-11 aircraft systems can be manually controlled from the overhead area of the cockpit. The center portion of the overhead panel is composed of the primary aircraft systems panels, which include FUEL, AIR, Electrical (ELEC) and Hydraulic (HYD) systems, which are easily accessible from both flight crew positions. Each Aircraft Systems Controller (ASC) has two automatic channels and a manual mode. All rectangular lights are annunciators. All square lights are combined switches and annunciators called switch/lights. Red switch/lights on the overhead (Level 3 alerts) are for conditions requiring immediate crew action. Amber (Level 2 or Level 1 alerts) indicates a fault or switch out of position requiring awareness or crew interaction. Overhead switches used in normal operating conditions will illuminate blue when in use (Level 0 alerts) such as WING ANTI-ICE - ON. An overhead switch/light with BLACK LETTERING on an amber or red background indicates a system failure and that crew interaction is required. A switch/light with blue or amber lettering and a BLACK BACKGROUND indicates a switch out of normal position and that crew action is necessary only if the system is in manual operation.

Oldale, Paul↗
NASA NTRS1952

Calculation of Internal Pressures in the Fuel Tube of a Nuclear Reactor

General procedures for computing internal pressures in fuel tubes of nuclear reactors are described and the effects on the pressure of varying neutron flux, fissioning material, and operating temperatures are discussed. A general proof is given that during pile operation each fission product is monotonically increasing and therefore a maximum amount of all elements is present at the time of shit down. The post-shutdown build-up of elements that are held in check during pile operation because of their inordinately high capture cross sections is calculated quantitatively. An account of chemical interactions between the many fission-product elements and the resulting effect on the total pressure completes the discussion. The general methods are illustrated by calculations applied to a system consisting of 90 percent enriched U235 in the form of UO2 packed into a hollow metal cylinder or "pin", operating at a flux of 8 x 10(exp 14) at 2000 F. Calculations of the pressure inside a pin are made with and without a sodium metal heat-transfer additive. The bulk of the pressure is shown to depend on the four elements, xenon, krypton, rubidium, and cesium; the amount of free oxygen, however, was also significant. For a shutdown time of 10(exp 6) seconds, the pressure was about 100 atmospheres.

Rosenbaum, B. M.↗
NASA NTRS1993

STS-55 pad abort: Engine 2011 oxidizer preburner augmented spark igniter check valve leak

The STS-55 initial launch attempt of Columbia (OV102) was terminated on KSC launch pad A March 22, 1993 at 9:51 AM E.S.T. due to violation of an ME-3 (Engine 2011) Launch Commit Criteria (LCC) limit exceedance. The event description and timeline are summarized. Propellant loading was initiated on 22 March, 1993 at 1:15 AM EST. All SSME chill parameters and launch commit criteria (LCC) were nominal. At engine start plus 1.44 seconds, a Failure Identification (FID) was posted against Engine 2011 for exceeding the 50 psia Oxidizer Preburner (OPB) purge pressure redline. The engine was shut down at 1.50 seconds followed by Engines 2034 and 2030. All shut down sequences were nominal and the mission was safely aborted. The OPB purge pressure redline violation and the abort profile/overlay for all three engines are depicted. SSME Avionics hardware and software performed nominally during the incident. A review of vehicle data table (VDT) data and controller software logic revealed no failure indications other than the single FID 013-414, OPB purge pressure redline exceeded. Software logic was executed according to requirements and there was no anomalous controller software operation. Immediately following the abort, a Rocketdyne/NASA failure investigation team was assembled. The team successfully isolated the failure cause to the oxidizer preburner augmented spark igniter purge check valve not being fully closed due to contamination. The source of the contaminant was traced to a cut segment from a rubber O-ring which was used in a fine clean tool during valve production prior to 1992. The valve was apparently contaminated during its fabrication in 1985. The valve had performed acceptably on four previous flights of the engine, and SSME flight history shows 780 combined check valve flights without failure. The failure of an Engine 3 (SSME No. 2011) check valve to close was sensed by onboard engine instruments even though all other engine operations were normal. This resulted in an engine shutdown and safe sequential shutdown of all three engines prior to ignition of the solid boosters.

Source record↗
NASA NTRS1993

STS-51 pad abort. OV103-engine 2033 (ME-2) fuel flowmeter sensor open circuit

The STS-51 initial launch attempt of Discovery (OV-103) was terminated on KSC launch pad 39B on 12 Aug. 1993 at 9:12 AM E.S.T. due to a sensor redundancy failure in the liquid hydrogen system of ME-2 (Engine 2033). The event description and time line are summarized. Propellant loading was initiated on 12 Aug. 1993 at 12:00 AM EST. All space shuttle main engine (SSME) chill parameters and Launch Commit Criteria (LCC) were nominal. At engine start plus 1.34 seconds a Failure Identification (FID) was posted against Engine 2033 for exceeding the 1800 spin intra-channel (A1-A2) Fuel Flowrate sensor channel qualification limit. The engine was shut down at 1.50 seconds followed by Engines 2032 and 2030. All shut down sequences were nominal and the mission was safely aborted. SSME Avionics hardware and software performed nominally during the incident. A review of vehicle data table (VDT) data and controller software logic revealed no failure indications other than the single FID 111-101, Fuel Flowrate Intra-Channel Test Channel A disqualification. Software logic was executed according to requirements and there was no anomalous controller software operation. Immediately following the abort, a Rocketdyne/NASA failure investigation team was assembled. The team successfully isolated the failure cause to an open circuit in a Fuel Flowrate Sensor. This type of failure has occurred eight previous times in ground testing. The sensor had performed acceptably on three previous flights of the engine and SSME flight history shows 684 combined fuel flow rate sensor channel flights without failure. The disqualification of an Engine 2 (SSME No. 2033) Fuel Flowrate sensor channel was a result of an instrumentation failure and not engine performance. All other engine operations were nominal. This disqualification resulted in an engine shutdown and safe sequential shutdown of all three engines prior to ignition of the solid boosters.

Source record↗
NASA NTRS1994

Mutagenesis in human cells with accelerated H and Fe ions

The overall goals of this research were to determine the risks of mutation induction and the spectra of mutations induced by energetic protons and iron ions at two loci in human lymphoid cells. During the three year grant period the research has focused in three major areas: (1) the acquisition of sufficient statistics for human TK6 cell mutation experiments using Fe ions (400 MeV/amu), Fe ions (600 MeV/amu) and protons (250 MeV/amu); (2) collection of thymidine kinase- deficient (tk) mutants or hypoxanthine phosphoribosyltransferase deficient (hprt) mutants induced by either Fe 400 MeV/amu, Fe 600 MeV/amu, or H 250 MeV/amu for subsequent molecular analysis; and (3) molecular characterization of mutants isolated after exposure to Fe ions (600 MeV/amu). As a result of the shutdown of the BEVALAC heavy ion accelerator in December 1992, efforts were rearranged somewhat in time to complete our dose-response studies and to complete mutant collections in particular for the Fe ion beams prior to the shutdown. These goals have been achieved. A major effort was placed on collection, re-screening, and archiving of 3 different series of mutants for the various particle beam exposures: tk-ng mutants, tk-sg mutants, and hprt-deficient mutants. Where possible, groups of mutants were isolated for several particle fluences. Comparative analysis of mutation spectra has occured with characterization of the mutation spectrum for hprt-deficient mutants obtained after exposure of TK6 cells to Fe ions (600 MeV/amu) and a series of spontaneous mutants.

Kronenberg, Amy↗
NASA NTRS1995

Display-And-Alarm Circuit For Accelerometer

Compact accelerometer assembly consists of commercial accelerometer retrofit with display-and-alarm circuit. Provides simple means for technician attending machine to monitor vibrations. Also simpifies automatic safety shutdown by providing local alarm or shutdown signal when vibration exceeds preset level.

Bozeman, Richard J., Jr.↗
NASA NTRS1996

Ampoule Failure System

An ampoule failure system for use in material processing furnaces comprising a containment cartridge and an ampoule failure sensor. The containment cartridge contains an ampoule of toxic material therein and is positioned within a furnace for processing. An ampoule failure probe is positioned in the containment cartridge adjacent the ampoule for detecting a potential harmful release of toxic material therefrom during processing. The failure probe is spaced a predetermined distance from the ampoule and is chemically chosen so as to undergo a timely chemical reaction with the toxic material upon the harmful release thereof. The ampoule failure system further comprises a data acquisition system which is positioned externally of the furnace and is electrically connected to the ampoule failure probe so as to form a communicating electrical circuit. The data acquisition system includes an automatic shutdown device for shutting down the furnace upon the harmful release of toxic material. It also includes a resistance measuring device for measuring the resistance of the failure probe during processing. The chemical reaction causes a step increase in resistance of the failure probe whereupon the automatic shutdown device will responsively shut down the furnace.

Watring, Dale A.↗
NASA NTRS1998

The Control System for the X-33 Linear Aerospike Engine

The linear aerospike engine is being developed for single-stage -to-orbit (SSTO) applications. The primary advantages of a linear aerospike engine over a conventional bell nozzle engine include altitude compensation, which provides enhanced performance, and lower vehicle weight resulting from the integration of the engine into the vehicle structure. A feature of this integration is the ability to provide thrust vector control (TVC) by differential throttling of the engine combustion elements, rather than the more conventional approach of gimballing the entire engine. An analysis of the X-33 flight trajectories has shown that it is necessary to provide +/- 15% roll, pitch and yaw TVC authority with an optional capability of +/- 30% pitch at select times during the mission. The TVC performance requirements for X-33 engine became a major driver in the design of the engine control system. The thrust level of the X-33 engine as well as the amount of TVC are managed by a control system which consists of electronic, instrumentation, propellant valves, electro-mechanical actuators, spark igniters, and harnesses. The engine control system is responsible for the thrust control, mixture ratio control, thrust vector control, engine health monitoring, and communication to the vehicle during all operational modes of the engine (checkout, pre-start, start, main-stage, shutdown and post shutdown). The methodology for thrust vector control, the health monitoring approach which includes failure detection, isolation, and response, and the basic control system design are the topic of this paper. As an additional point of interest a brief description of the X-33 engine system will be included in this paper.

Jackson, Jerry E.↗