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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 NTRS1971

A study of LC-39 cryogenic systems. Part 1: A study of the vacuum insulated transfer lines at Kennedy Space Center. Part 2: Cooldown pressure surges

The vacuum liquid hydrogen and liquid oxygen transfer lines at Kennedy Space Center were studied to evaluate the feasibility of using a condensing gas such as CO2 inside the vacuum spaces to achieve a condensing-vacuum. The study indicates that at ambient temperature, a maximum vacuum hyphen space pressure of 4000 microns is acceptable for the LH2 transfer lines. In addition, the cooldown procedures for the 14-inch cross-country liquid oxygen line was studied using a simplified mathematical model. Preliminary cooldown times are presented for various heat leak rates to the line and for two vent configurations.

Ludtke, P. R.↗
NASA NTRS2019

Orion Capsule Parachute Assembly System (CPAS) Overload Testing Approach and Results

The qualification of the Orion Capsule Parachute Assembly System (CPAS) system includes exposure to loads and dynamic pressures above the required values as validation of the parachutes' structural integrity. As outlined in the certification plan, each of the four parachutes of the system are to be subjected to 110% of their respective maximum dynamic pressure requirements. The Main and Drogue parachutes have satisfied this overload condition in drop testing and due to cost and schedule constraints, the Forward Bay Cover Parachute (FBCP) and Pilot parachute were subjected to the overload condition in the ground testing described in this document. The test objectives and pass/fail criteria were established and require the parachutes to achieve and maintain a target riser load (associated with a minimum of 110% dynamic pressure overload) for a minimum of three seconds while sustaining no failures of any structural members (vent hoop, radials, suspension line) or any damage which propagates into catastrophic failure of the canopy. Considering the assumptions and limitations associated with the ground testing (primarily non-uniform flow field of the ground test system and variations in parachute manufacturing), a method of establishing the desired overload condition was determined by the technical community and covers the 2-sigma bounds of the drag area distribution derived from drop testing. On June 27, 2017 the testing was executed at the High Velocity Airflow System (HiVAS) facility located at the Weapon Survivability Laboratory (WSL) at the Naval Air Warfare Center, Weapons Division (NAWCWD) China Lake. Engineering Development Unit (EDU) parachutes were used as pathfinders to gain experience with achieving the test conditions. Additional runs were executed to measure the airflow at the same location as the canopy skirt, although this data is not required to satisfy the test objectives and parachute pass/fail criteria. The qualification parachutes were successfully exposed to the target conditions and sustained only minor damage.

Daum, Jared S.↗
NASA NTRS2011

Cryogenic Propellant Feed System Analytical Tool Development

The Propulsion Systems Branch at NASA s Lyndon B. Johnson Space Center (JSC) has developed a parametric analytical tool to address the need to rapidly predict heat leak into propellant distribution lines based on insulation type, installation technique, line supports, penetrations, and instrumentation. The Propellant Feed System Analytical Tool (PFSAT) will also determine the optimum orifice diameter for an optional thermodynamic vent system (TVS) to counteract heat leak into the feed line and ensure temperature constraints at the end of the feed line are met. PFSAT was developed primarily using Fortran 90 code because of its number crunching power and the capability to directly access real fluid property subroutines in the Reference Fluid Thermodynamic and Transport Properties (REFPROP) Database developed by NIST. A Microsoft Excel front end user interface was implemented to provide convenient portability of PFSAT among a wide variety of potential users and its ability to utilize a user-friendly graphical user interface (GUI) developed in Visual Basic for Applications (VBA). The focus of PFSAT is on-orbit reaction control systems and orbital maneuvering systems, but it may be used to predict heat leak into ground-based transfer lines as well. PFSAT is expected to be used for rapid initial design of cryogenic propellant distribution lines and thermodynamic vent systems. Once validated, PFSAT will support concept trades for a variety of cryogenic fluid transfer systems on spacecraft, including planetary landers, transfer vehicles, and propellant depots, as well as surface-based transfer systems. The details of the development of PFSAT, its user interface, and the program structure will be presented.

Lusby, Brian S.↗
NASA NTRS1986

Basaltic fissure eruptions, plume heights, and atmospheric aerosols

Convective plumes that rise above Hawaiian-style fire fountains consist of volcanic gases, aerosols, fine ash, and entrained heated air. Plume theory has been applied to observational estimates of the rate of thermal energy release from large fire fountains. The theoretically predicted heights of maintained plumes agree very well with the heights found from actual observations. Predicted plume heights for both central-vent (point-source) and fissure (line-source) eruptions indicate a stratospheric penetration by plumes that form over vents with very high magma-production rates. Flood basalt fissure eruptions that produce individual lava flows with volumes greater than 100 cu km at very high mass eruption rates are capable of injecting large quantities of sulfate aerosols into the lower stratosphere, with potentially drastic short-term atmospheric consequences, like acid precipitation, darkening of the sky, and climatic cooling.

Stothers, R. B.↗
NASA NTRS2013

Temperature Stratification in a Cryogenic Fuel Tank

A reduced dynamical model describing temperature stratification effects driven by natural convection in a liquid hydrogen cryogenic fuel tank has been developed. It accounts for cryogenic propellant loading, storage, and unloading in the conditions of normal, increased, and micro- gravity. The model involves multiple horizontal control volumes in both liquid and ullage spaces. Temperature and velocity boundary layers at the tank walls are taken into account by using correlation relations. Heat exchange involving the tank wall is considered by means of the lumped-parameter method. By employing basic conservation laws, the model takes into consideration the major multi-phase mass and energy exchange processes involved, such as condensation-evaporation of the hydrogen, as well as flows of hydrogen liquid and vapor in the presence of pressurizing helium gas. The model involves a liquid hydrogen feed line and a tank ullage vent valve for pressure control. The temperature stratification effects are investigated, including in the presence of vent valve oscillations. A simulation of temperature stratification effects in a generic cryogenic tank has been implemented in Matlab and results are presented for various tank conditions.

stratification↗
NASA NTRS2015

CFD Simulations of the Space Launch System Ascent Aerodynamics and Booster Separation

This paper presents details of Computational Fluid Dynamic modeling of the Space Launch System during ascent. The primary focus of the paper is the flow simulation of the vehicle during ascent using the Overflow Navier-Stokes code. Computations of 739 first-stage flight conditions covering a range of Mach numbers, angles of attack, and roll angles were computed. The overset grid system contained 375 million grid points, and over 28 mil- lion CPU hours were used in the simulations. The simulations were run on the Pleiades supercomputer at the NASA Advanced Supercomputer Center at Ames Research Center. The data products from this work include integrated line-loads, surface pressure coefficients, venting pressures, and protuberance air-loads. Detailed comparisons were made of the aerodynamic performance predicted by Overflow and the wind-tunnel derived aero- dynamic database. A small number of the cases were run with two different turbulence models and with two differencing schemes. These results were used to quantify the sensitivity to the choice of the turbulence model and to the differencing scheme. The paper also introduces an effort to use the inviscid, unstructured Cartesian solver Cart3D to compute the aerodynamics during booster separation. Adaptive mesh refinement is being used to enable accurate simulations of sixteen booster-separation-motor plumes. The use of this tool is explored in preparation for building a booster-separation aerodynamic database.

Ascent Aerodynamics↗
NASA NTRS1990

Cryogenic On-Orbit Liquid Depot Storage, Acquisition, and Transfer Satellite (COLD-SAT)

The Cryogenic On-Orbit Liquid Depot Storage, Acquisition, and Transfer Satellite (COLD-SAT) will perform subcritical liquid hydrogen handling experiments under low gravity conditions to provide engineering data for future space transportation missions. Comprising the four Class 1 enabling experiments are tank press control, tank chilldown, tank no-vent fill, and liquid acquisition device fill/refill. The nine Class 2 enhancing experiments are tanker thermal performance, pressurization, low-gravity setting and outflow, liquid acquisition device performance, transfer line chilldown, outflow subcooling, low-gravity vented fill, fluid dumping, and advanced instrumentation. Consisting of an experiment module mated to a spacecraft bus, COLD-SAT will be placed in an initial 1300 km circular orbit by an Atlas commercial launch vehicle, and will perform experiments in a semi-autonomous mode for a period of up to six months. The three-axis controlled spacecraft bus provides electric power, control and data management, communications, and attitude control along with propulsive acceleration levels ranging from 10(exp -6) to 10(exp -4) g. It is desired to understand the effects that low acceleration levels might have on the heat and mass transfer processes involved in some of the experiments. The experiment module contains the three liquid hydrogen tanks, valves, pressurization and pumping equipment, and instrumentation. Within the highly insulated tanks are specialized fluid management equipment that might be used in future space transportation systems. At launch all the liquid hydrogen for the experiments is contained in the largest tank, which has helium-purged insulation to prevent cryo-pumping of air on the launch pad. The tank is loaded by the hydrogen tanking system used for the Centaur upper stage of the Atlas. After reaching orbit the two smaller tanks become receivers for fluid transfers, and when tanked, become the vessels for performing many of the experiments.

Schuster, John R.↗
NASA NTRS2005

Making AlN(x) Tunnel Barriers Using a Low-Energy Nitrogen-Ion Beam

A technique based on accelerating positive nitrogen ions onto an aluminum layer has been demonstrated to be effective in forming thin (<2 nm thick) layers of aluminum nitride (AlN(x)) for use as tunnel barriers in Nb/Al-AlN(x)/Nb superconductor/insulator/ superconductor (SIS) Josephson junctions. AlN(x) is the present material of choice for tunnel barriers because, to a degree greater than that of any other suitable material, it offers the required combination of low leakage current at high current density and greater thermal stability. While ultra-thin AlN films with good thickness and stoichiometry control are easily formed using techniques such as reactive molecular beam epitaxy and chemical vapor deposition, growth temperatures of 900 C are necessary for the dissociative adsorption of nitrogen from either nitrogen (N2) or ammonia (NH3). These growth temperatures are prohibitively high for the formation of tunnel barriers on Nb films because interfacial reactions at temperatures as low as 200 to 300 C degrade device properties. Heretofore, deposition by reactive sputtering and nitridation of thin Al layers with DC and RF nitrogen plasmas have been successfully used to form AlN barriers in SIS junctions. However, precise control over critical current density Jc has proven to be a challenge, as is attaining adequate process reproducibility from system to system. The present ion-beam technique is an alternative to the plasma or reactive sputtering techniques as it provides a highly controlled arrival of reactive species, independent of the electrical conditions of the substrate or vacuum chamber. Independent and accurate control of parameters such as ion energy, flux, species, and direction promises more precise control of film characteristics such as stoichiometry and thickness than is the case with typical plasma processes. In particular, the background pressure during ion-beam nitride growth is 2 or 3 orders of magnitude lower, minimizing the formation of compounds with contaminants, which is critical in devices the performance of which is dictated by interfacial characteristics. In addition, the flux of incoming species can be measured in situ using ion probes so that the dose can be controlled accurately. The apparatus used in the present ion-beam technique includes a vacuum chamber containing a commercial collimated- ion-beam source, a supply of nitrogen and argon, and an ion probe for measuring the ion dose. Either argon or nitrogen can be used as the feed gases for the ion source, depending on whether cleaning of the substrate or growth of the nitride, respectively, is desired. Once the Nb base electrode and Al proximity layer have been deposited, the N2 gas line to the ion beam is vented and purged, and the ion-source is turned on until a stable discharge is obtained. The substrate is moved over the ion-beam source to expose the Al surface layer to the ion beam (see figure) for a specified duration for the formation of the nitride tunnel barrier. Next, the Nb counter-electrode layer is deposited on the nitride surface layer. The Nb/Al- AlN(x)/Nb-trilayer-covered substrate is then patterned into individual devices by use of conventional integrated-circuit processing techniques.

Kaul, Anupama↗
NASA NTRS1993

Design and integrated operation of an innovative thermodynamic vent system concept

A unique zero-g thermodynamic vent system (TVS) is being developed by NASA's Marshall Space Flight Center (MSFC) and Rockwell International to meet cryogenic propellant management requirements for future space missions. The design is highly innovative in that it integrates the functions of a spray-bar tank mixer and a TVS. This concept not only satisfies the requirement for efficient tank mixing and zero-g venting but also accommodates thermal conditioning requirements for other components (e.g., engine feed lines, turbopumps, and liquid acquisition devices). In addition, operations can be extended to accomplish tank chill-down, no-vent fill, and emergency venting during zero-g propellant transfer. This paper describes the system performance characterization and future test activities that are part of MSFC's Multipurpose Hydrogen Test Bed (MHTB) program. The testing will demonstrate the feasibility and merit of the design, and serve as a proof-of-concept development activity.

Fazah, Michel M.↗
NASA NTRS2016

Vented Chill / No-Vent Fill of Cryogenic Propellant Tanks

Architectures for extended duration missions often include an on-orbit replenishment of the space vehicle's cryogenic liquid propellants. Such a replenishment could be accomplished via a tank-to-tank transfer from a dedicated tanker or a more permanent propellant depot storage tank. Minimizing the propellant loss associated with transfer line and receiver propellant tank thermal conditioning is essential for mass savings. A new methodology for conducting tank-to-tank transfer while minimizing such losses has been demonstrated. Charge-Hold-Vent is the traditional methodology for conducting a tank-to-tank propellant transfer. A small amount of cryogenic liquid is introduced to chill the transfer line and propellant tank. As the propellant absorbs heat and undergoes a phase change, the tank internal pressure increases. The tank is then vented to relieve pressure prior to another charge of cryogenic liquid being introduced. This cycle is repeated until the transfer lines and tank are sufficiently chilled and the replenishment of the propellant tank is complete. This method suffers inefficiencies due to multiple chill and vent cycles within the transfer lines and associated feed system components. Additionally, this system requires precise measuring of cryogenic fluid delivery for each transfer, multiple valve cycling events, and other complexities associated with cycled operations. To minimize propellant loss and greatly simplify on-orbit operations, an alternate methodology has been designed and demonstrated. The Vented Chill / No Vent Fill method is a simpler, constant flow approach in which the propellant tank and transfer lines are only chilled once. The receiver tank is continuously vented as cryogenic liquid chills the transfer lines, tank mass and ullage space. Once chilled sufficiently, the receiver tank valve is closed and the tank is completely filled. Interestingly, the vent valve can be closed prior to receiver tank components reaching liquid saturation temperature. An incomplete fill results if insufficient energy is removed from the tank's thermal mass and ullage space. The key to successfully conducting the no vent fill is to assure that sufficient energy is removed from the system prior to closing the receiver tank vent valve. This paper will provide a description of the transfer methodology and test article, and will provide a discussion of test results.

Rhys, Noah O.↗
NASA NTRS1977

Critical velocity phenomena and the LTP

When the relative velocity between magnetized plasma and neutral gas exceeds a critical value, the gas-plasma interaction is dominated by collective phenomena which rapidly excite and ionize the neutrals. The interaction of the solar wind with a large cloud (between 10 to the 24th and 10 to the 28th power neutrals) vented from the moon should be of this type. Line radiation from such an interaction can yield an apparent lunar surface brightness rivaling reflected sunlight levels over small areas, if the kinetic-energy flow density of the gas is sufficiently high. The aberrated solar-wind flow past the moon would enhance the visibility of such interactions near the lunar sunrise terminator, supporting the statistical studies which indicate that the 'Lunar Transient Phenomena' (anomalous optical phenomena on the moon) are significantly correlated with the position of the terminator on the lunar surface.

Srnka, L. J.↗
NASA NTRS1974

Cryogenic thermal control technology summaries

A summarization and categorization is presented of the pertinent literature associated with cryogenic thermal control technology having potential application to in-orbit fluid transfer systems and/or associated space storage. Initially, a literature search was conducted to obtain pertinent documents for review. Reports determined to be of primary significance were summarized in detail. Each summary, where applicable, consists of; (1) report identification, (2) objective(s) of the work, (3) description of pertinent work performed, (4)major results, and (5) comments of the reviewer (GD/C). Specific areas covered are; (1) multilayer insulation of storage tanks with and without vacuum jacketing, (2) other insulation such as foams, shadow shields, microspheres, honeycomb, vent cooling and composites, (3) vacuum jacketed and composite fluid lines, and (4) low conductive tank supports and insulation penetrations. Reports which were reviewed and not summarized, along with reasons for not summarizing, are also listed.

Stark, J. A.↗
NASA NTRS2007

Low Velocity Airdrop Tests of an X-38 Backup Parachute Design

The NASA Johnson Space Center's X-38 program designed a new backup parachute system to recover the 25,000 lb X-38 prototype for the Crew Return Vehicle spacecraft. Due to weight and cost constraints, the main backup parachute design incorporated rapid and low cost fabrication techniques using off-the-shelf materials. Near the vent, the canopy was constructed of continuous ribbons, to provide more damage tolerance. The remainder of the canopy was a constructed with a continuous ringslot design. After cancellation of the X-38 program, the parachute design was resized, built, and drop tested for Natick Soldiers Center's Low Velocity Air Drop (LVAD) program to deliver cargo loads up to 22,000 lbs from altitudes as low as 500 feet above the ground. Drop tests results showed that the 500-foot LVAD parachute deployment conditions cause severe skirt inversion and inflation problems for large parachutes. The bag strip occurred at a high angle of attack, causing skirt inversion before the parachute could inflate. The addition of a short reefing line prevented the skirt inversion. Using a lower porosity in the vent area, than is normally used in large parachutes, improved inflation. The drop testing demonstrated that the parachute design could be refined to meet the requirements for the 500-foot LVAD mission.

Stein, Jenny M.↗
NASA NTRS1990

Flow fields of low pressure vent exhausts

The flow field produced by low pressure gas vents are described based on experimental data obtained from tests in a large vacuum chamber. The gas density, pressure, and flux at any location in the flow field are calculated based on the vent plume description and the knowledge of the flow rate and velocity of the venting gas. The same parameters and the column densities along a specified line of sight traversing the plume are also obtained and shown by a computer generated graphical representation. The fields obtained with a radically scanning Pitot probe within the exhausting gas are described by a power of the cosine function, the mass rate, and the distance from the exit port. The field measurements were made for gas at pressures ranging from 2 to 50 torr venting from pipe fittings with diameters to 3/16 to 1-1/2 inches I.D. (4.76 to 38.1 mm). The N2 mass flow rates ranged from 2E-4 to 3.7E-1 g/s.

Scialdone, John J.↗
NASA NTRS1989

Flow fields of low pressure vent exhausts

The flow field produced by low pressure gas vents are described based on experimental data obtained from tests in a large vacuum chamber. The gas density, pressure, and flux at any location in the flow field are calculated based on the vent plume description and the knowledge of the flow rate and velocity of the venting gas. The same parameters and the column densities along a specified line of sight traversing the plume are also obtained and shown by a computer-generated graphical representation. The fields obtained with a radially scanning Pitot probe within the exhausting gas are described by a power of the cosine function, the mass rate and the distance from the exit port. The field measurements were made for gas at pressures ranging from 2 to 50 torr venting from pipe fittings with diameters of 3/16 inch to 1-1/2 inches I.D. (4.76 mm to 38.1 mm). The N(2) mass flow rates ranged from 2E-4 to 3.7E-1 g/s.

Scialdone, John J.↗
NASA NTRS2006

Testing the Effects of Helium Pressurant on Thermodynamic Vent System Performance with Liquid Hydrogen

In support of the development of a zero gravity pressure control capability for liquid hydrogen, testing was conducted at the Marshall Space Flight Center using the Multipurpose Hydrogen Test Bed (MHTB) to evaluate the effects of helium pressurant on the performance of a spray bar thermodynamic vent system (TVS). Fourteen days of testing was performed in August - September 2005, with an ambient heat leak of about 70-80 watts and tank fill levels of 90%, 50%, and 25%. The TVS successfully controlled the tank pressure within a +/- 3.45 kPa (+/- 0.5 psi) band with various helium concentration levels in the ullage. Relative to pressure control with an "all hydrogen" ullage, the helium presence resulted in 10 to 30 per cent longer pressure reduction durations, depending on the fill level, during the mixing/venting phase of the control cycle. Additionally, the automated control cycle was based on mixing alone for pressure reduction until the pressure versus time slope became positive, at which time the Joule-Thomson vent was opened. Testing was also conducted to evaluate thermodynamic venting without the mixer operating, first with liquid then with vapor at the recirculation line inlet. Although ullage stratification was present, the ullage pressure was successfully controlled without the mixer operating. Thus, if vapor surrounded the pump inlet in a reduced gravity situation, the ullage pressure can still be controlled by venting through the TVS Joule Thomson valve and heat exchanger. It was evident that the spray bar configuration, which extends almost the entire length of the tank, enabled significant thermal energy removal from the ullage even without the mixer operating. Details regarding the test setup and procedures are presented in the paper. 1

Flachbart, R. H.↗
NASA NTRS2017

Pressure-Fed LOX/LCH4 Reaction Control System for Spacecraft: Transient Modeling and Thermal Vacuum Hotfire Test Results

An integrated cryogenic liquid oxygen, liquid methane (LOX/LCH4) reaction control system (RCS) was tested at NASA Glenn Research Center's Plum Brook Station in the Spacecraft Propulsion Research Facility (B-2) under vacuum and thermal vacuum conditions. The RCS is a subsystem of the Integrated Cryogenic Propulsion Test Article (ICPTA), a pressure-fed LOX/LCH4 propulsion system composed of a single 2,800 lbf main engine, two 28 lbf RCS engines, and two 7 lbf RCS engines. Propellants are stored in four 48 inch diameter 5083 aluminum tanks that feed both the main engine and RCS engines in parallel. Helium stored cryogenically in a composite overwrapped pressure vessel (COPV) flows through a heat exchanger on the main engine before being used to pressurize the propellant tanks to a design operating pressure of 325 psi. The ICPTA is capable of simultaneous main engine and RCS operation. The RCS engines utilize a coil-on-plug (COP) ignition system designed for operation in a vacuum environment, eliminating corona discharge issues associated with a high voltage lead. There are two RCS pods on the ICPTA, with two engines on each pod. One of these two engines is a heritage flight engine from Project Morpheus. Its sea level nozzle was removed and replaced by an 85:1 nozzle machined using Inconel 718, resulting in a maximum thrust of 28 lbf under altitude conditions. The other engine is a scaled down version of the 28 lbf engine, designed to match the core and overall mixture ratios as well as other injector characteristics. This engine can produce a maximum thrust of 7 lbf with an 85:1 nozzle that was additively manufactured using Inconel 718. Both engines are film-cooled and capable of limited duration gas-gas and gas-liquid operation, as well as steady-state liquid-liquid operation. Each pod contains one of each version, such that two engines of the same thrust level can be fired as a couple on opposite pods. The RCS feed system is composed of symmetrical 3/8 inch lines that tap off of the main propellant manifold to send LOX and LCH4 outboard to the RCS pods. A Thermodynamic Vent System (TVS) is used to condition propellants at each pod by venting through an orifice and then routing the cold expansion products back through tubing that is welded along a large portion of the main RCS feed lines. Prior to final installation on the ICPTA, the RCS engines were tested in a small vacuum chamber at the Johnson Space Center (JSC) Energy Systems Test Area (ESTA) to verify functionality of the new COP ignition system and check out operation of the vacuum nozzles. After engine-level testing, the RCS engines were installed on the vehicle and a series of integrated hot-fire tests were performed at JSC consisting of various pulsing and steady-state firings as well as integrated main engine/RCS operation. The ICPTA was then integrated into the Plum Brook B-2 facility for vacuum and thermal/vacuum testing. Testing in the B-2 facility was composed of multiple thermal and pressure environments. The first set of tests were performed under ambient temperature and altitude pressure conditions. These tests consisted of a range of minimum impulse bit (MIB) pulsing sequences with low duty cycle, analogous to a coast phase in which the RCS is primarily used for station keeping. The primary goal of this sequence is to understand how propellant conditions were effected without an active TVS. In this scenario, consistent gas-gas operation is desirable since it results in a smaller MIB and more efficient propellant consumption. Multiple skin thermocouples are mounted on the feedlines, in addition to a submerged thermocouple on each commodity, in order to gather thermal data on the system. Higher duty cycle pulsing tests were then performed, analogous to an ascent or landing mission phase. The primary goal of this sequence was to examine how well the engines self-conditioned without active TVS when starting from a quiescent state. The TVS was then activated during some tests to demonstrate the capability to quickly condition the engines for higher pulsing demand scenarios. A thermocouple at the TVS outlet allows for the calculation of energy absorbed by the vented propellant. Lastly, tests with longer pulses and multiple engines firing either in sequence or simultaneously were run in order to gather transient system response data on waterhammer. Six total high-speed pressure transducers are installed on the RCS system, one sensor at the end of each propellant manifold line on the pods, and one at the tap-off location for each commodity. This will allow for the accurate characterization of waterhammer in the system under various propellant conditions and firing sequences. Other instrumentation for this test series includes nozzle throat thermocouples, chamber pressure measurement, heat soakback measurement, and tank wall plume impingement temperature measurement. The next set of tests were performed to demonstrate simultaneous main engine and RCS operation. Data from this test will be used to examine if there is any change to nominal operation of the RCS as a result of feed system interaction or other phenomenon. Some of these tests began under high vacuum conditions (target ambient pressure less than 1x10(exp -3) torr) and others began at altitude conditions. The last set of tests were performed with the B-2 cold wall active. Under these tests, many of the same low duty cycle MIB tests were repeated in order to characterize how propellant conditions changed with the lower heat leak. In this scenario the RCS manifold experiences much less heat leak, resulting in a change to how well the engines self-condition. As a result, an increase in maximum waterhammer pressures and a change in natural frequency of the system was expected due to higher density propellants. The lower heat leak should also result in a change to the MIB pulse profile, and data will be examined to understand how MIB repeatability is affected in the different operating environments. Parallel to the test efforts, a set of transient model development efforts were made to predict RCS performance. The primary effort was aimed at producing a SINDA/FLUINT model to predict propellant conditioning up to the engine inlet as a function of different environmental and operating parameters, with the goal of predicting chamber pressure, TVS performance, and propellant consumption over time. Preliminary results for this effort will be presented in comparison with test data. Additional modeling efforts were made using SINDA/FLUINT to predict waterhammer in the system since the software is capable of handling multiphase transient fluid dynamics. These results will be compared with the high-speed pressure transducer test data for validation purposes.

Atwell, Matthew J.↗
NASA NTRS1980

Wind tunnel investigation of the Titan Forward Skirt compartment vent from a free-stream Mach number of 0.80 to 1.96

A test was conducted to determine the flow characteristics of the Titan forward skirt compartment vent over a free stream Mach number range of 0.80 to 1.96. The vent was mounted in a flat plate and the plate was flush mounted to the tunnel side wall with coinciding center lines. Air was discharged from a duct, located on the tunnel side wall behind the plate, through a canted aft 30 deg honeycomb vent into the free stream. Data for the analysis of the Titan forward skirt compartment venting during ascent through the atmosphere are provided. Full scale simulated flight hardware, such as the honeycomb vent, duct corrugations and field joint ring were used. Boundary layer thicknesses were used to vary boundary height. The highest vent discharge coefficient for any given Mach number and vent pressure ratio generally occurred at the maximum displacement thickness. With no vent flow the static pressure in the vent region was generally less than the free stream static pressure. With vent flow, the static pressures upstream of the vent increased, and those downstream of the vent decreased.

Johns, A. L.↗