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NASA NTRS2008

Calibrating the Helium Pressurization System for the Space Shuttle Liquid-Hydrogen Tank

Analysis of the results from the STS-114 tanking tests and subsequent launch called into question existing thermal and mass models of helium pressurization of the liquid hydrogen tank. This hydrogen tank, which makes up the bottom two-thirds of the External Tank, is pressurized prior to launch to avoid cavitation in the Shuttle Main Engine pumps. At about 2 minutes prior to launch, the main vent valve is closed, and pressurized helium flows into the tank ullage space to achieve set point pressure. As the helium gas cools, its pressure drops, calling for additional helium. Subsequent helium flows are provided in short, timed pulses. The number of pulses is taken as a rough leak indicator. An analysis of thermal models by Marshall Space Flight Center showed considerable uncertainty in the pressure-versus-time behavior of the helium ullage space and the ability to predict the number of pulses normally expected. Kennedy Space Center proposed to calibrate the dime-sized orifice, which together with valves, controls the helium flow quantity (Figure 1). Pressure and temperature sensors were installed to provide upstream and downstream measurements necessary to compute flow rate based on the orifice discharge coefficient. An assessment of flow testing with helium indicated an extremely costly use of this critical resource. In order to reduce costs, we proposed removing the orifices from each Mobile Launcher Platform (MLP) and asking Colorado Engineering Experiment Station Inc. (CEESI) to calibrate the flow. CEESI has a high-pressure air flow system with traceable flow meters capable of handling the large flow rates. However, literature research indicated that square-edged orifices of small diameters often exhibit significant hysteresis and nonrepeatability in the vicinity of choked or sonic flow. Fortunately, the MLP orifices behaved relatively well in testing (Figure 2). Using curve fitting of the air-flow data, in conjunction with ASME orifice modeling equations, a method of relating the helium mass flow to measured air flow data was obtained. This analysis showed that the highest uncertainty in flow occurred in the vicinity of the choking pressure ratio, as would be expected. In addition, analysis of typical flow pulses showed that most of the helium flow occurred either well below or well above this uncertain area. The final result is the ability to provide postlaunch estimates of helium mass flows that are within 1.5 percent of the actual value.

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NASA NTRS1949

Calibration Tests of a German Log Rodmeter

A German log rodmeter of the pitot static type was calibrated in Langley tank no. 1 at speeds up to 34 knots and angles of yaw from 0 deg to plus or minus 10 3/4 degrees. The dynamic head approximated the theoretical head at 0 degrees yaw but decreased as the yaw was increased. The static head was negative and in general became more negative with increasing speed and yaw. Cavitation occurred at speeds above 31 knots at 0 deg yaw and 21 knots at 10 3/4 deg yaw.

Mottard, Elmo J.↗

NASA NTRS1949

Calibration Tests of a Japanese Log Rodmeter

A Japanese log rodmeter of the rotating-vane impeller type, with a co mmutator on the impeller shaft, was calibrated in Langley tank no. 1. The rotational speed of two impellers was determined for forward speeds up to 24 knots at angles of yaw up to ?10 0 . In general, the rota tional speeds of two apparently identical impellers tested in the rod meter decreased with increasing yaw angle, right yaw causing a greate r decrease than left yaw. The difference in calibration between the t wo impellers was approximately the same as that produced by a change in yaw angle from 50 left to 50 right. Evidence of cavitation within the impeller fairing appeared at speeds above 24 knots.

Mottard, Elmo J.↗

NASA NTRS2007

Analyzing the Use of Gaseous Helium as a Pressurant with Cryogenic Propellants with Thermodynamic Venting System Modelling and Test Data

Cryogens are viable candidate propellants for NASA's Lunar and Mars exploration programs. To provide adequate mass flow to the system's engines and/or to prevent feed system cavitation, gaseous helium (GHe) is frequently considered as a pressurant. During low gravity operations, a Thermodynamic Venting System (TVS) is designed to maintain tank pressure during low gravity operations without propellant resettling. Therefore, a series of tests were conducted in the Multi-purpose Hydrogen Test Bed (MHTB) of Marshall Space Flight Center (MSFC) in order to evaluate the effects of GHe pressurant on pressure control performance of a TVS with liquid hydrogen (LH2) and nitrogen (LN2) as the test liquids. The TVS used in these test series consists of a recirculation pump, Joule-Thomson (J-T) expansion valve, and a parallel flow concentric tube heat exchanger combined with a longitudinal spray bar. Using a small amount of liquid extracted from the tank recirculation line, passing it through the J-T valve, and then through the heat exchanger, thermal energy is extracted from the bulk liquid and ullage thereby enabling pressure control. The LH2/GHe tests were performed at fill levels of 90%, 50%, and 25% and LN2/GHe tests were conducted at fill levels of 50% and 25%. Moreover, each test was conducted with a specified tank ullage pressure control band. A one-dimensional TVS performance program was used to analyze and correlate the test data. Predictions and comparisons with test data of ullage pressure and temperature and bulk liquid saturation pressure and temperature with test data are presented.

Hedayat, A.↗

NASA NTRS2007

Cryogenic Capillary Screen Heat Entrapment

Cryogenic liquid acquisition devices (LADs) for space-based propulsion interface directly with the feed system, which can be a significant heat leak source. Further, the accumulation of thermal energy within LAD channels can lead to the loss of sub-cooled propellant conditions and result in feed system cavitation during propellant outflow. Therefore, the fundamental question addressed by this program was: "To what degree is natural convection in a cryogenic liquid constrained by the capillary screen meshes envisioned for LADs.?"Testing was first conducted with water as the test fluid, followed by LN2 tests. In either case, the basic experimental approach was to heat the bottom of a cylindrical column of test fluid to establish stratification patterns measured by temperature sensors located above and below a horizontal screen barrier position. Experimentation was performed without barriers, with screens, and with a solid barrier. The two screen meshes tested were those typically used by LAD designers, "200x1400" and "325x2300", both with Twill Dutch Weave. Upon consideration of both the water and LN2 data it was concluded that heat transfer across the screen meshes was dependent upon barrier thermal conductivity and that the capillary screen meshes were impervious to natural convection currents.

Bolshinskiy, L.G.↗

NASA NTRS2007

NASA Tech Briefs, October 2007

Topics covered include; Wirelessly Interrogated Position or Displacement Sensors; Ka-Band Radar Terminal Descent Sensor; Metal/Metal Oxide Differential Electrode pH Sensors; Improved Sensing Coils for SQUIDs; Inductive Linear-Position Sensor/Limit-Sensor Units; Hilbert-Curve Fractal Antenna With Radiation- Pattern Diversity; Single-Camera Panoramic-Imaging Systems; Interface Electronic Circuitry for an Electronic Tongue; Inexpensive Clock for Displaying Planetary or Sidereal Time; Efficient Switching Arrangement for (N + 1)/N Redundancy; Lightweight Reflectarray Antenna for 7.115 and 32 GHz; Opto-Electronic Oscillator Using Suppressed Phase Modulation; Alternative Controller for a Fiber-Optic Switch; Strong, Lightweight, Porous Materials; Nanowicks; Lightweight Thermal Protection System for Atmospheric Entry; Rapid and Quiet Drill; Hydrogen Peroxide Concentrator; MMIC Amplifiers for 90 to 130 GHz; Robot Would Climb Steep Terrain; Measuring Dynamic Transfer Functions of Cavitating Pumps; Advanced Resistive Exercise Device; Rapid Engineering of Three-Dimensional, Multicellular Tissues With Polymeric Scaffolds; Resonant Tunneling Spin Pump; Enhancing Spin Filters by Use of Bulk Inversion Asymmetry; Optical Magnetometer Incorporating Photonic Crystals; WGM-Resonator/Tapered-Waveguide White-Light Sensor Optics; Raman-Suppressing Coupling for Optical Parametric Oscillator; CO2-Reduction Primary Cell for Use on Venus; Cold Atom Source Containing Multiple Magneto- Optical Traps; POD Model Reconstruction for Gray-Box Fault Detection; System for Estimating Horizontal Velocity During Descent; Software Framework for Peer Data-Management Services; Autogen Version 2.0; Tracking-Data-Conversion Tool; NASA Enterprise Visual Analysis; Advanced Reference Counting Pointers for Better Performance; C Namelist Facility; and Efficient Mosaicking of Spitzer Space Telescope Images.

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NASA NTRS2010

Minnowbrook VI: 2009 Workshop on Flow Physics and Control for Internal and External Aerodynamics

Topics covered include: Flow Physics and control for Internal and External Aerodynamics (not in TOC...starts on pg13); Breaking CFD Bottlenecks in Gas-Turbine Flow-Path Design; Streamwise Vortices on the Convex Surfaces of Circular Cylinders and Turbomachinery Blading; DNS and Embedded DNS as Tools for Investigating Unsteady Heat Transfer Phenomena in Turbines; Cavitation, Flow Structure and Turbulence in the Tip Region of a Rotor Blade; Development and Application of Plasma Actuators for Active Control of High-Speed and High Reynolds Number Flows; Active Flow Control of Lifting Surface With Flap-Current Activities and Future Directions; Closed-Loop Control of Vortex Formation in Separated Flows; Global Instability on Laminar Separation Bubbles-Revisited; Very Large-Scale Motions in Smooth and Rough Wall Boundary Layers; Instability of a Supersonic Boundary-Layer With Localized Roughness; Active Control of Open Cavities; Amplitude Scaling of Active Separation Control; U.S. Air Force Research Laboratory's Need for Flow Physics and Control With Applications Involving Aero-Optics and Weapon Bay Cavities; Some Issues Related to Integrating Active Flow Control With Flight Control; Active Flow Control Strategies Using Surface Pressure Measurements; Reduction of Unsteady Forcing in a Vaned, Contra-Rotating Transonic Turbine Configuration; Active Flow Control Stator With Coanda Surface; Controlling Separation in Turbomachines; Flow Control on Low-Pressure Turbine Airfoils Using Vortex Generator Jets; Reduced Order Modeling Incompressible Flows; Study and Control of Flow Past Disk, and Circular and Rectangular Cylinders Aligned in the Flow; Periodic Forcing of a Turbulent Axisymmetric Wake; Control of Vortex Breakdown in Critical Swirl Regime Using Azimuthal Forcing; External and Turbomachinery Flow Control Working Group; Boundary Layers, Transitions and Separation; Efficiency Considerations in Low Pressure Turbines; Summary of Conference; and Final Plenary Session Transcript.

LaGraff, John E.↗

NASA NTRS2010

SSME Investment in Turbomachinery Inducer Impeller Design Tools and Methodology

Within the rocket engine industry, SSME turbomachines are the de facto standards of success with regard to meeting aggressive performance requirements under challenging operational environments. Over the Shuttle era, SSME has invested heavily in our national inducer impeller design infrastructure. While both low and high pressure turbopump failures/anomaly resolution efforts spurred some of these investments, the SSME program was a major benefactor of key areas of turbomachinery inducer-impeller research outside of flight manifest pressures. Over the past several decades, key turbopump internal environments have been interrogated via highly instrumented hot-fire and cold-flow testing. Likewise, SSME has sponsored the advancement of time accurate and cavitating inducer impeller computation fluid dynamics (CFD) tools. These investments together have led to a better understanding of the complex internal flow fields within aggressive high performing inducers and impellers. New design tools and methodologies have evolved which intend to provide confident blade designs which strike an appropriate balance between performance and self induced load management.

Zoladz, Thomas↗

NASA NTRS2010

A Historical Systems Study of Liquid Rocket Engine Throttling Capabilities

This is a comprehensive systems study to examine and evaluate throttling capabilities of liquid rocket engines. The focus of this study is on engine components, and how the interactions of these components are considered for throttling applications. First, an assessment of space mission requirements is performed to determine what applications require engine throttling. A background on liquid rocket engine throttling is provided, along with the basic equations that are used to predict performance. Three engines are discussed that have successfully demonstrated throttling. Next, the engine system is broken down into components to discuss special considerations that need to be made for engine throttling. This study focuses on liquid rocket engines that have demonstrated operational capability on American space launch vehicles, starting with the Apollo vehicle engines and ending with current technology demonstrations. Both deep throttling and shallow throttling engines are discussed. Boost and sustainer engines have demonstrated throttling from 17% to 100% thrust, while upper stage and lunar lander engines have demonstrated throttling in excess of 10% to 100% thrust. The key difficulty in throttling liquid rocket engines is maintaining an adequate pressure drop across the injector, which is necessary to provide propellant atomization and mixing. For the combustion chamber, cooling can be an issue at low thrust levels. For turbomachinery, the primary considerations are to avoid cavitation, stall, surge, and to consider bearing leakage flows, rotordynamics, and structural dynamics. For valves, it is necessary to design valves and actuators that can achieve accurate flow control at all thrust levels. It is also important to assess the amount of nozzle flow separation that can be tolerated at low thrust levels for ground testing.

Betts, Erin M.↗

NASA NTRS2010

The Evolution of Utilizing Manual Throttling to Avoid Excessively Low LH2 NPSP at the SSME Inlet

In the late 1970s, years before the Space Shuttle flew its maiden voyage, it was understood low liquid hydrogen (LH2) Net Positive Suction Pressure (NPSP) at the inlet to the Space Shuttle Main Engine (SSME) could have adverse effects on engine operation. A number of failures within both the External Tank (ET) and the Orbiter Main Propulsion System (MPS) could result in a low LH2 NPSP condition, which at extremely low levels can result in cavitation of SSME turbomachinery. Operational workarounds were developed to take advantage of the onboard crew s ability to manually throttle down the SSMEs (via the Pilot s Speedbrake/Throttle Controller), which alleviated the low LH2 NPSP condition. Manually throttling the SSME to a lower power level resulted in an increase in NPSP, mainly due to the reduction in frictional flow losses while at the lower throttle setting. Early in the Space Shuttle Program s history, the relevant Flight Rule for the Booster flight controller in Mission Control did not distinguish between ET and Orbiter MPS failures and the same crew action was taken for both. However, after a review of all Booster operational techniques following the Challenger disaster in the late 1980s, it was determined manually throttling the SSME to a lower power was only effective for Orbiter MPS failures and the Flight Rule was updated to reflect this change. The Flight Rule and associated crew actions initially called for a single throttle step to minimum power level when a low threshold for NPSP was met. As engineers refined their understanding of the NPSP requirements for the SSME (through a robust testing program), the operational techniques evolved to take advantage of the additional capabilities. This paper will examine the evolution of the Flight rule and associated procedure and how increases in knowledge about the SSME and the Space Shuttle vehicle as a whole have helped shape their development. What once was a single throttle step when NPSP decreased to a certain low threshold has now become three throttle steps, each occurring at a lower NPSP threshold. Additionally the procedure, which for early Space Shuttle missions required a Return-to-Launch-Site abort, now results in a nominal Main Engine Cut Off and no loss of mission objectives.

Henfling, Rick↗

NASA NTRS2011

Hazards Induced by Breach of Liquid Rocket Fuel Tanks: Conditions and Risks of Cryogenic Liquid Hydrogen-Oxygen Mixture Explosions

We analyze the data of purposeful rupture experiments with LOx and LH2 tanks, the Hydrogen-Oxygen Vertical Impact (HOVI) tests that were performed to clarify the ignition mechanisms, the explosive power of cryogenic H2/Ox mixtures under different conditions, and to elucidate the puzzling source of the initial formation of flames near the intertank section during the Challenger disaster. We carry out a physics-based analysis of general explosions scenarios for cryogenic gaseous H2/Ox mixtures and determine their realizability conditions, using the well-established simplified models from the detonation and deflagration theory. We study the features of aerosol H2/Ox mixture combustion and show, in particular, that aerosols intensify the deflagration flames and can induce detonation for any ignition mechanism. We propose a cavitation-induced mechanism of self-ignition of cryogenic H2/Ox mixtures that may be realized when gaseous H2 and Ox flows are mixed with a liquid Ox turbulent stream, as occurred in all HOVI tests. We present an overview of the HOVI tests to make conclusion on the risk of strong explosions in possible liquid rocket incidents and provide a semi-quantitative interpretation of the HOVI data based on aerosol combustion. We uncover the most dangerous situations and discuss the foreseeable risks which can arise in space missions and lead to tragic outcomes. Our analysis relates to only unconfined mixtures that are likely to arise as a result of liquid propellant space vehicle incidents.

Osipov, Viatcheslav↗

NASA NTRS2006

System for Continuous Deaeration of Hydraulic Oil

A system for continuous, rapid deaeration of hydraulic oil has been built to replace a prior system that effected deaeration more slowly in a cyclic pressure/ vacuum process. Such systems are needed because (1) hydraulic oil has an affinity for air, typically containing between 10 and 15 volume percent of air and (2) in the original application for which these systems were built, there is a requirement to keep the proportion of dissolved air below 1 volume percent because a greater proportion can lead to pump cavitation and excessive softness in hydraulic-actuator force-versus-displacement characteristics. In addition to overcoming several deficiencies of the prior deaeration system, the present system removes water from the oil. The system (see figure) includes a pump that continuously circulates oil at a rate of 10 gal/min (38 L/min) between an 80-gal (303-L) airless reservoir and a tank containing a vacuum. When the circulation pump is started, oil is pumped, at a pressure of 120 psi (827 kPa), through a venturi tube below the tank with a connection to a stand-pipe in the tank. This action draws oil out of the tank via the standpipe. At the same time, oil is sprayed into the tank in a fine mist, thereby exposing a large amount of oil to the vacuum. When the oil level in the tank falls below the lower of two level switches, a vacuum pump is started, drawing a hard vacuum on the tank through a trap that collects any oil and water entrained in the airflow. When the oil level rises above higher of the two level switches or when the system is shut down, a solenoid valve between the tank and the vacuum pump is closed to prevent suction of oil into the vacuum pump. Critical requirements that the system is designed to satisfy include the following: 1) The circulation pump must have sufficient volume and pressure to operate the venturi tube and spray nozzles. 2) The venturi tube must be sized to empty the tank (except for the oil retained by the standpipe) and maintain a vacuum against the vacuum pump. 3) The tank must be strong enough to withstand atmospheric pressure against the vacuum inside and must have sufficient volume to enable exposure of a sufficiently large amount of sprayed oil to the vacuum. 4) The spray nozzles must be sized to atomize the oil and to ensure that the rate of flow of sprayed oil does not exceed the rate at which the venturi action can empty the tank. 5) The vacuum pump must produce a hard vacuum against the venturi tube and continue to work when it ingests some oil and water. 6) Fittings must be made vacuum tight (by use of O-rings) to prevent leakage of air into the system. The system is fully automatic, and can be allowed to remain in operation with very little monitoring. It is capable of reducing the air content of the oil from 11 to less than 1 volume percent in about 4 hours and to keep the water content below 100 parts per million.

Anderson, Christopher W.↗

NASA NTRS2011

The Evolution of Utilizing Manual Throttles to Avoid Excessively Low LH2 NPSP at the SSME Inlet

In the late 1970s, years before the Space Shuttle flew its maiden voyage, it was understood low liquid hydrogen (LH2) Net Positive Suction Pressure (NPSP) at the inlet to the Space Shuttle Main Engine (SSME) could have adverse effects on engine operation. A number of failures within both the External Tank (ET) and the Orbiter Main Propulsion System (MPS) could result in a low LH2 NPSP condition, which at extremely low levels can result in cavitation of SSME turbomachinery. Operational workarounds were developed to take advantage of the onboard crew s ability to manually throttle down the SSMEs (via the Pilot s Speedbrake/Throttle Controller), which alleviated the low LH2 NPSP condition. Manually throttling the SSME to a lower power level resulted in an increase in NPSP, mainly due to the reduction in frictional flow losses while at the lower throttle setting. Early in the Space Shuttle Program s history, the relevant Flight Rule for the Booster flight controllers in Mission Control did not distinguish between ET and Orbiter MPS failures and the same crew action was taken for both. However, after a review of all Booster operational techniques following the Challenger disaster in the late 1980s, it was determined manually throttling the SSME to a lower power was only effective for Orbiter MPS failures and the Flight Rule was updated to reflect this change. The Flight Rule and associated crew actions initially called for a single throttle step to minimum power level when a low threshold for NPSP was met. As engineers refined their understanding of the NPSP requirements for the SSME (through a robust testing program), the operational techniques evolved to take advantage of the additional capabilities. This paper will examine the evolution of the Flight rule and associated procedure and how increases in knowledge about the SSME and the Space Shuttle vehicle as a whole have helped shape their development. What once was a single throttle step when NPSP decreased to a certain threshold has now become three throttle steps, each occurring at a lower NPSP threshold. Additionally the procedure, which for early Space Shuttle missions required a Return-to-Launch-Site abort, now results in a nominal Main Engine Cut Off and no loss of mission objectives.

Henfling, Rick↗

NASA NTRS2005

NASA Tech Briefs, November 2005

Topics covered include: Laser System for Precise, Unambiguous Range Measurements; Flexible Cryogenic Temperature and Liquid-Level Probes; Precision Cryogenic Dilatometer; Stroboscopic Interferometer for Measuring Mirror Vibrations; Some Improvements in H-PDLCs; Multiple-Bit Differential Detection of OQPSK; Absolute Position Encoders With Vertical Image Binning; Flexible, Carbon-Based Ohmic Contacts for Organic Transistors; GaAs QWIP Array Containing More Than a Million Pixels; AutoChem; Virtual Machine Language; Two-Dimensional Ffowcs Williams/Hawkings Equation Solver; Full Multigrid Flow Solver; Doclet To Synthesize UML; Computing Thermal Effects of Cavitation in Cryogenic Liquids; GUI for Computational Simulation of a Propellant Mixer; Control Program for an Optical-Calibration Robot; SQL-RAMS; Distributing Data from Desktop to Hand-Held Computers; Best-Fit Conic Approximation of Spacecraft Trajectory; Improved Charge-Transfer Fluorescent Dyes; Stability-Augmentation Devices for Miniature Aircraft; Tool Measures Depths of Defects on a Case Tang Joint; Two Heat-Transfer Improvements for Gas Liquefiers; Controlling Force and Depth in Friction Stir Welding; Spill-Resistant Alkali-Metal-Vapor Dispenser; A Methodology for Quantifying Certain Design Requirements During the Design Phase; Measuring Two Key Parameters of H3 Color Centers in Diamond; Improved Compression of Wavelet-Transformed Images; NASA Interactive Forms Type Interface - NIFTI; Predicting Numbers of Problems in Development of Software; Hot-Electron Photon Counters for Detecting Terahertz Photons; Magnetic Variations Associated With Solar Flares; and Artificial Intelligence for Controlling Robotic Aircraft.

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NASA NTRS2011

In-Service Monitoring of Steam Pipe Systems at High Temperatures

An effective, in-service health monitoring system is needed to track water condensation in real time through the walls of steam pipes. The system is required to measure the height of the condensed water from outside the pipe, while operating at temperatures that are as high as 250 C. The system needs to account for the effects of water flow and cavitation. In addition, it is desired that the system does not require perforating the pipes and thereby reducing the structural integrity. Generally, steam pipes are used as part of the district heating system carrying steam from central power stations under the streets to heat, cool, or supply power to high-rise buildings and businesses. This system uses ultrasonic waves in pulse-echo and acquires reflected signal data. Via autocorrelation, it determines the water height while eliminating the effect of noise and multiple reflections from the wall of the pipe. The system performs nondestructive monitoring through the walls of steam pipes, and automatically measures the height of condensed water while operating at the high-temperature conditions of 250 C. For this purpose, the ultrasonic pulse-echo method is used where the time-of-flight of the wave reflections inside the water are measured, and it is multiplied by the wave velocity to determine the height. The pulse-echo test consists of emitting ultrasonic wave pulses from a piezoelectric transducer and receiving the reflections from the top and bottom of the condensed water. A single transducer is used as a transmitter as well as the receiver of the ultrasonic waves. To obtain high resolution, a broadband transducer is used and the frequency can be in the range of 2.25 to 10 MHz, providing sharp pulses in the time domain allowing for higher resolution in identifying the individual reflections.

Bar-Cohen, Yoseph↗

NASA NTRS2011

Experimental Determination of the Dynamic Hydraulic Transfer Function for the J-2X Oxidizer Turbopump: Methodology - Part One

An advanced methodology for extracting the hydraulic dynamic pump transfer matrix (Yp) for a cavitating liquid rocket engine turbopump inducer+impeller has been developed. The transfer function is required for integrated vehicle pogo stability analysis as well as optimization of local inducer pumping stability. Laboratory pulsed subscale waterflow test of the J-2X oxygen turbo pump is introduced and our new extraction method applied to the data collected. From accurate measures of pump inlet and discharge perturbational mass flows and pressures, and one-dimensional flow models that represents complete waterflow loop physics, we are able to derive Yp and hence extract the characteristic pump parameters: compliance, pump gain, impedance, mass flow gain. Detailed modeling is necessary to accurately translate instrument plane measurements to the pump inlet and discharge and extract Yp. We present the MSFC Dynamic Lump Parameter Fluid Model Framework and describe critical dynamic component details. We report on fit minimization techniques, cost (fitness) function derivation, and resulting model fits to our experimental data are presented. Comparisons are made to alternate techniques for spatially translating measurement stations to actual pump inlet and discharge.

Zoladz, Tom↗

NASA NTRS2011

High Temperature Monitoring the Height of Condensed Water in Steam Pipes

An in-service health monitoring system is needed for steam pipes to track through their wall the condensation of water. The system is required to measure the height of the condensed water inside the pipe while operating at temperatures that are as high as 250 deg. C. The system needs to be able to make real time measurements while accounting for the effects of cavitation and wavy water surface. For this purpose, ultrasonic wave in pulse-echo configuration was used and reflected signals were acquired and auto-correlated to remove noise from the data and determine the water height. Transmitting and receiving the waves is done by piezoelectric transducers having Curie temperature that is significantly higher than 250 deg. C. Measurements were made at temperatures as high as 250 deg. C and have shown the feasibility of the test method. This manuscript reports the results of this feasibility study.

fluid height monitoring↗

NASA NTRS2013

Powerpack Testing of the J2-X Oxidizer Turbopump

The J-2X PPA-2 (Powerpack 2) test series was conducted from February through December 2012 on the A1 test stand at NASA Stennis Space Center in a joint effort between Pratt and Whitney Rocketdyne (PWR) and NASA Marshall Space Flight Center (MSFC). The series consisted of 13 tests for a total hotfire duration close to 6500 seconds. Among the chief test objectives was characterization of the capabilities of the new J-2X turbopumps. This paper concentrates on the test results pertaining specifically to the Oxidizer Turbopump (OTP) operation in liquid oxygen (LOX). The two bladed inducer configuration that was tested had a 0.81% blade height tip clearance and incorporated the B-groove modification to the inducer tunnel outboard of the inducer leading edge. Data was collected on inducer suction performance, pump radial loading, and the pump dynamic environment in LOX. Comparisons were made to prior data collected on two geometrically similar subscale inducers tested in water at NASA MSFC (70% scale) and Concepts NREC (52% scale). In overview, the results of the powerpack testing were consistent with the radial load assessment from the water tests. The inducer performed differently in the other two categories, however. The inducer suction performance capability in LOX was notably improved. This was expected and is probably attributable to thermal suppression head (TSH) effects. While the dynamic environment was similar in most aspects to water test observations, the higher order cavitation (HOC) during the powerpack testing was much more benign.

Jambusaria, Mitul↗