Results for “Propellant Tank” — Page 86

NASA NTRS1990

The COLD-SAT Experiment for Cryogenic Fluid Management Technology

Future national space transportation missions will depend on the use of cryogenic fluid management technology development needs for these missions. In-space testing will be conducted in order to show low gravity cryogenic fluid management concepts and to acquire a technical data base. Liquid H2 is the preferred test fluid due to its propellant use. The design of COLD-SAT (Cryogenic On-orbit Liquid Depot Storage, Acquisition, and Transfer Satellite), an Expendable Launch Vehicle (ELV) launched orbital spacecraft that will perform subcritical liquid H2 storage and transfer experiments under low gravity conditions is studied. An Atlas launch vehicle will place COLD-SAT into a circular orbit, and the 3-axis controlled spacecraft bus will provide electric power, experiment control, and data management, attitude control, and propulsive accelerations for the experiments. Low levels of acceleration will provide data on the effects that low gravity might have on the heat and mass transfer processes used. The experiment module will contain 3 liquid H2 tanks; fluid transfer, pressurization and venting equipment; and instrumentation.

Schuster, J. R.↗

NASA NTRS2017

Multilayer Insulation Ascent Venting Model

The thermal and venting transient experienced by tank-applied multilayer insulation (MLI) in the Earth-to-orbit environment is very dynamic and not well characterized. This new predictive code is a first principles-based engineering model which tracks the time history of the mass and temperature (internal energy) of the gas in each MLI layer. A continuum-based model is used for early portions of the trajectory while a kinetic theory-based model is used for the later portions of the trajectory, and the models are blended based on a reference mean free path. This new capability should improve understanding of the Earth-to-orbit transient and enable better insulation system designs for in-space cryogenic propellant systems.

Tramel, R. W.↗

NASA NTRS2019

Versatile Nuclear Thermal Propulsion (NTP)

A study was initiated to investigate propulsion stage and mission architecture options potentially enabled by fission energy. One initial concept is a versatile Nuclear Thermal Propulsion (NTP) system with a maximum specific impulse of 900 s and a maximum thrust (per engine) of 15 klbf. The system assumes a monopropellant stage (hydrogen), and is designed to also provide 300 lbf of thrust (potentially split between multiple thrusters) at an Isp > 500 s for orbital maneuvering and station keeping. Boost pumps are used to assist with engine decay heat removal and low thrust engine burns, and to compensate for partial tank depressurization during full thrust engine burns. Potential stage assembly orbits that take full advantage of launch vehicle payload mass and volume capabilities are being assessed. The potential for using NTP engines to also generate a small to moderate amount of electrical power is also being evaluated. A first generation versatile NTP stage could enable 8 of 9 upcoming opportunities for short (less than 24 month) round trip human missions to Mars. A second generation versatile NTP is under consideration that could potentially provide a maximum specific impulse of 1800 s at 15 klbf, and enable ambitious missions throughout the solar system. The second generation NTP system under consideration would also allow a choice of volatiles to be used as propellant. This would potentially allow in-situ resources such as water, ammonia, methane, or other compounds to be used directly as propellant by the second generation engine.

Houts, Michael G.↗

NASA NTRS2019

Versatile Nuclear Thermal Propulsion (NTP)

A study was initiated to investigate propulsion stage and mission architecture options potentially enabled by fission energy. One initial concept is a versatile Nuclear Thermal Propulsion (NTP) system with a maximum specific impulse of 900 s and a maximum thrust (per engine) of 15 klbf. The system assumes a monopropellant stage (hydrogen), and is designed to also provide 300 lbf of thrust (potentially split between multiple thrusters) at an Isp > 500 s for orbital maneuvering and station keeping. Boost pumps are used to assist with engine decay heat removal and low thrust engine burns, and to compensate for partial tank depressurization during full thrust engine burns. Potential stage assembly orbits that take full advantage of launch vehicle payload mass and volume capabilities are being assessed. The potential for using NTP engines to also generate a small to moderate amount of electrical power is also being evaluated. A first generation versatile NTP stage could enable 8 of 9 upcoming opportunities for short (less than 24 month) round trip human missions to Mars. A second generation versatile NTP is under consideration that could potentially provide a maximum specific impulse of 1800 s at 15 klbf, and enable ambitious missions throughout the solar system. The second generation NTP system under consideration would also allow a choice of volatiles to be used as propellant. This would potentially allow in-situ resources such as water, ammonia, methane, or other compounds to be used directly as propellant by the second generation engine.

Houts, Mike↗

NASA NTRS2000

Magnetically Actuated Propellant Orientation Experiment, Controlling Fluid Motion With Magnetic Fields in a Low-Gravity Environment

This report details the results of a series of fluid motion experiments to investigate the use of magnets to orient fluids in a low-gravity environment. The fluid of interest for this project was liquid oxygen (LO2) since it exhibits a paramagnetic behavior (is attracted to magnetic fields). However, due to safety and handling concerns, a water-based ferromagnetic mixture (produced by Ferrofluidics Corporation) was selected to simplify procedures. Three ferromagnetic fluid mixture strengths and a nonmagnetic water baseline were tested using three different initial fluid positions with respect to the magnet. Experiment accelerometer data were used with a modified computational fluid dynamics code termed CFX-4 (by AEA Technologies) to predict fluid motion. These predictions compared favorably with experiment video data, verifying the code's ability to predict fluid motion with and without magnetic influences. Additional predictions were generated for LO2 with the same test conditions and geometries used in the testing. Test hardware consisted of a cylindrical Plexiglas tank (6-in. bore with 10-in. length), a 6,000-G rare Earth magnet (10-in. ring), three-axis accelerometer package, and a video recorder system. All tests were conducted aboard the NASA Reduced-Gravity Workshop, a KC-135A aircraft.

Martin, J. J.↗

NASA NTRS2017

Large Scale Production of Densified Hydrogen Using Integrated Refrigeration and Storage

Recent demonstration of advanced liquid hydrogen storage techniques using Integrated Refrigeration and Storage (IRAS) technology at NASA Kennedy Space Center led to the production of large quantities of solid densified liquid and slush hydrogen in a 125,000 L tank. Production of densified hydrogen was performed at three different liquid levels and LH2 temperatures were measured by twenty silicon diode temperature sensors. System energy balances and solid mass fractions are calculated. Experimental data reveal hydrogen temperatures dropped well below the triple point during testing (up to 1 K), and were continuing to trend downward prior to system shutdown. Sub-triple point temperatures were seen to evolve in a time dependent manner along the length of the horizontal, cylindrical vessel. Twenty silicon diode temperature sensors were recorded over approximately one month for testing at two different fill levels (33 67). The phenomenon, observed at both two fill levels, is described and presented detailed and explained herein., and The implications of using IRAS for energy storage, propellant densification, and future cryofuel systems are discussed.

Notardonato, William U.↗

NASA NTRS1947

Hydrodynamic Qualities of a 1/10-Size Powered Dynamic Model of the XP5Y-1 Flying Boat in Smooth Water: Langley Tank Model 246, TED No. NACA DE 320

The hydrodynamic characteristics of a 1/10-size powered dynamic model of the XP5Y-1 flying boat were determined in Langley tank no. 1. Stable take-offs were possible at all practicable positions of the center of gravity and flap deflections. An increase in gross load from 123.5 to 150.0 pounds (21.5 percent) had only a slight effect on the stable range for take-off. A decrease in forward acceleration from 3.0 to 1.0 feet per second per second had only a very small effect on the stable range for take-off. In general, the landings were free from skipping except at trims below 6 deg where one skip was encountered at an aft position of the center of gravity. The model porpoised during the landing runout at all positions of the center of gravity when landed at trims above 10 deg. Spray in the propellers was light at the design gross load, and was not considered excessive,at a gross load of 136.0 pounds.

Woodward, David R.↗

NASA NTRS2004

One Idea for a Next Generation Shuttle

In this configuration, the current Shuttle External Tank serves as core structure for a fully reusable second stage. This stage is equipped with wings, vertical fin, landing gear, and thermal protection. The stage is geometrically identical to (but smaller than) a single stage that has been tested hyper-sonically, super-sonically, and sub-sonically in the NASA Langley Research Center wind tunnels. The three LOX/LH engines that currently serve as main propulsion for the Shuttle Orbiter, serve as main propulsion on the new stage. The new stage is unmanned but is equipped with the avionics needed for automatic maneuvering on orbit and for landing on a runway. Three rails are installed along the top surface of the vehicle for attachment of various payloads. Pay- loads might include third stages with satellites attached, personnel pods, propellants, or other items.

MacConochie, Ian O.↗

NASA NTRS2000

Magnetically Actuated Propellant Orientation, Controlling Fluids in a Low-Gravity Environment

Cryogenic fluid management (CFM) is a technology area common to virtually every space transportation propulsion concept envisioned. Storage, supply, transfer and handling of sub-critical cryogenic fluids are basic capabilities that have long been needed by multiple programs and the need is expected to continue in the future. The use of magnetic fields provides another method, which could replace or augment current/traditional approaches, potentially simplifying vehicle operational constraints. The magnetically actuated propellant orientation (MAPO) program effort focused on the use of magnetic fields to control fluid motion as it relates to positioning (i.e. orientation and acquisition) of a paramagnetic substance such as LO2. Current CFM state- of-the-art systems used to control and acquire propellant in low gravity environments rely on liquid surface tension devices which employ vanes, fine screen mesh channels and baskets. These devices trap and direct propellant to areas where it's needed and have been used routinely with storable (non-cryogenic) propellants. However, almost no data exists r,egarding their operation in cryogenics and the use of such devices confronts designers with a multitude of significant technology issues. Typical problems include a sensitivity to screen dry out (due to thermal loads and pressurant gas) and momentary adverse accelerations (generated from either internal or external sources). Any of these problems can potentially cause the acquisition systems to ingest or develop vapor and fail. The use of lightweight high field strength magnets may offer a valuable means of augmenting traditional systems potentially mitigating or at least easing operational requirements. Two potential uses of magnetic fields include: 1) strategically positioning magnets to keep vent ports clear of liquid (enabling low G vented fill operations), and 2) placing magnets in the center or around the walls of the tank to create an insulating vapor pocket (between the liquid and the tank wall) which could effectively lower heat transfer to the liquid (enabling increased storage time).

Martin, James J.↗

NASA NTRS2012

High-Pressure Systems Suppress Fires in Seconds

Much deserved attention is given to the feats of innovation that allow humans to live in space and robotic explorers to beam never-before-seen images back to Earth. In the background of these accomplishments is a technology that makes it all possible the rockets that propel NASA s space exploration efforts skyward. Marshall Space Flight Center has been at the heart of the Agency s rocketry and spacecraft propulsion efforts since its founding in 1960. Located at the Redstone Arsenal near Huntsville, Alabama, the Center has a legacy of success stretching back to the Saturn rockets that carried the Apollo astronauts into space. Even before Marshall was established, Redstone was the site of significant advances in American rocketry under the guidance of famous rocket engineer Werner Von Braun; these included the Juno I rocket that successfully carried the United States first satellite, Explorer 1, into orbit in 1958. And from the first orbital test flight of the Space Shuttle Columbia through the final flights of the shuttle program this year, these vehicles have been enabled by the solid rocket boosters, external tank, and orbiter main engines created at Marshall. Today, Marshall continues to host innovation in rocket and spacecraft propulsion at state-of-the-art facilities such as the Propulsion Research Laboratory. Like many of its past successes, some of the Center s current advancements are being made with the help of private industry partners. The efforts have led not only to new propulsion technologies, but to terrestrial benefits in a seemingly unrelated field in this case, firefighting.

Source record↗

NASA NTRS1974

Space shuttle SRM interim contract, part 1

Essential studies and analyses required to integrate the SRM into the booster and overall space shuttle system. Emphasis was placed on the case, nozzle, insulation, and propellant components with resulting performance, weight, and structural load characteristics being generated. Effort conducted during the time period of this contract included studies, analyses, planning, and preliminary design activities. Technical requirements identified in the SRM Project Request for Proposal No. 8-1-4-94-98401 and Thiokol's proposed SRM design (designated Configuration 0) established the basis for this effort. The requirements were evaluated jointly with MSFC and altered where necessary to incorporate new information that evolved after issuance of the RFP and during the course of this interim contract. Revised water impact loads and load distributions were provided based on additional model test data and analytical effort conducted by NASA subsequent to the RFP release. Launch pad peaking loads into the SRM aft skirt were provided which also represented a change from RFP requirements. A modified SRM/External Tank (ET) attachment configuration with new structural load data was supplied by NASA, and direction was received to include a 2 percent inert weight contingency.

Source record↗

NASA NTRS1966

Postflight Evaluation of Atlas-Centaur AC-6 (Launched August 11, 1965)

The sixth Atlas Centaur vehicle (AC-6) was successfully launched from the Eastern T e s t Range, Complex 36B, on August 11, 1965, at 0931:04.430 EST. A 2084-pound dynamic model of the Surveyor payload was placed in a simulated lunar transfer trajectory. Vehicle systems operated satisfactorily and all the flight objectives were accomplished. Lift-off within 4 seconds of the window opening demonstrated the launch-on-time capability of the vehicle were accurately compensated for by the Centaur guidance system. the Surveyor model into a near-perfect lunar transfer trajectory would have resulted in an impact of the moon without a midcourse correction. To hit the precise target area on the lunar surface, the required correction would have been 4.25 meters per second, which is well within the spacecraft capability. Normal thrust and impulse levels were obtained with both the A t l a s and Centaur propulsion systems. However, a sizeable thrust overshoot on startup of the Centaur engines has not been resolved. A propellant-utilization system used for the first time on the Centaur, accurately controlled the fuel and oxidant consumption. The turnaround and retrothrust maneuver were performed without incident. Relatively high longitudinal modal excitations and lateral payload excitations were obtained at lift-off; these high perturbations are believed t o be related t o the launcher holddown arms. Nominal temperatures were recorded for both the external vehicle skin and the payload compartment; however, abnormally low temperatures were measured in the forward equipment area, which may have resulted from leakage of cold helium purge gas. All vehicle electrical systems performed satisfactorily; the only difficulty with the RF systems was obtained with the C-band transponder. of the vehicle instrumentation yielded valid data. The AC-6 vehicle was constructed with several new lightweight designs including the forward bulkhead, thrust barrel, interstage adapter and tank skin thickness reduction from 0.016 t o 0.014 inch. No deficiencies were observed in any of these new structural elements.

Source record↗

NASA NTRS2015

Results of Microgravity Fluid Dynamics Captured With the Spheres-Slosh Experiment

This paper provides an overview of the SPHERES-Slosh Experiment (SSE) aboard the International Space Station (ISS) and presents on-orbit results with data analysis. In order to predict the location of the liquid propellant during all times of a spacecraft mission, engineers and mission analysts utilize Computational Fluid Dynamics (CFD). These state-of-the-art computer programs numerically solve the fluid flow equations to predict the location of the fluid at any point in time during different spacecraft maneuvers. The models and equations used by these programs have been extensively validated on the ground, but long duration data has never been acquired in a microgravity environment. The SSE aboard the ISS is designed to acquire this type of data, used by engineers on earth to validate and improve the CFD prediction models, improving the design of the next generation of space vehicles as well as the safety of current missions. The experiment makes use of two Synchronized Position Hold, Engage, Reorient Experimental Satellites (SPHERES) connected by a frame. In the center of the frame there is a plastic, pill shaped tank that is partially filled with green-colored water. A pair of high resolution cameras records the movement of the liquid inside the tank as the experiment maneuvers within the Japanese Experimental Module test volume. Inertial measurement units record the accelerations and rotations of the tank, making the combination of stereo imaging and inertial data the inputs for CFD model validation.

slosh↗

NASA NTRS2015

Results of Microgravity Fluid Dynamics Captured with the Spheres-Slosh Experiment

This paper provides an overview of the SPHERES-Slosh Experiment (SSE) aboard the International Space Station (ISS) and presents on-orbit results with data analysis. In order to predict the location of the liquid propellant during all times of a spacecraft mission, engineers and mission analysts utilize Computational Fluid Dynamics (CFD). These state-of-the-art computer programs numerically solve the fluid flow equations to predict the location of the fluid at any point in time during different spacecraft maneuvers. The models and equations used by these programs have been extensively validated on the ground, but long duration data has never been acquired in a microgravity environment. The SSE aboard the ISS is designed to acquire this type of data, used by engineers on earth to validate and improve the CFD prediction models, improving the design of the next generation of space vehicles as well as the safety of current missions. The experiment makes use of two Synchronized Position Hold, Engage, Reorient Experimental Satellites (SPHERES) connected by a frame. In the center of the frame there is a plastic, pill shaped tank that is partially filled with green-colored water. A pair of high resolution cameras records the movement of the liquid inside the tank as the experiment maneuvers within the Japanese Experimental Module test volume. Inertial measurement units record the accelerations and rotations of the tank, making the combination of stereo imaging and inertial data the inputs for CFD model validation.

slosh↗

NASA NTRS2015

Result of Microgravity Fluid Dynamics Captured with the SPHERES-Slosh Experiment

This paper provides an overview of the SPHERES-Slosh Experiment (SSE) aboard the International Space Station (ISS) and presents on-orbit results with data analysis. In order to predict the location of the liquid propellant during all times of a spacecraft mission, engineers and mission analysts utilize Computational Fluid Dynamics (CFD). These state-of-the-art computer programs numerically solve the fluid flow equations to predict the location of the fluid at any point in time during different spacecraft maneuvers. The models and equations used by these programs have been extensively validated on the ground, but long duration data has never been acquired in a microgravity environment. The SSE aboard the ISS is designed to acquire this type of data, used by engineers on earth to validate and improve the CFD prediction models, improving the design of the next generation of space vehicles as well as the safety of current missions. The experiment makes use of two Synchronized Position Hold, Engage, Reorient Experimental Satellites (SPHERES) connected by a frame. In the center of the frame there is a plastic, pill shaped tank that is partially filled with green-colored water. A pair of high resolution cameras records the movement of the liquid inside the tank as the experiment maneuvers within the Japanese Experimental Module test volume. Inertial measurement units record the accelerations and rotations of the tank, making the combination of stereo imaging and inertial data the inputs for CFD model validation.

slosh↗

NASA NTRS2004

Design of an EXB Probe

The amount of chemical propellant required to accomplish certain NASA s planned missions is too immense such that the spacecraft will never be able to lift off. To address this concern, electric propulsion systems have been chosen as the primary propulsion systems for some NASA s future missions, including DAWN and JIMO. Research Center is a proposed engine for the JIMO mission, which will visit three of Jupiter s icy moons. Optimizing thruster s lifetime and efficiency are the two foci for the engineers on the Ion Team. One qualitative study of the engine s efficiency can be accomplished by examining the ratio of doubly- to singly-charged ions in the ion beam of the engine. Thrust efficiency directly relates to this ratio. The bulk of this project is to redesign and build an EXB probe to obtain this qualitative measurement. Once this probe is built, it can be installed in a vacuum tank (VF 65 in building 301) behind the exit plane of the HiPEP engine to collect data. Current chemical propulsion technology cannot address the needs of some deep space The HiPEP (High Power Electric Propulsion) engine being developed at NASA Glenn Research Center is a proposed engine for the JIMO mission, which will visit three Jupiter's icy moons.

Nguyen, Sonca↗

NASA NTRS1999

Upper-Stage Flight Experiment

For propulsion applications that require that the propellants are storable for long periods, have a high density impulse, and are environmentally clean and non-toxic, the best choice is a combination of high-concentration hydrogen peroxide (High Test Peroxide, or HTP) and a liquid hydrocarbon (LHC) fuel. The HTP/LHC combination is suitable for low-cost launch vehicles, space taxi and space maneuvering vehicles, and kick stages. Orbital Sciences Corporation is under contract with the NASA Marshall Space Flight Center in cooperation with the Air Force Research Lab to design, develop and demonstrate a new low-cost liquid upper stage based on HTP and JP-8. The Upper Stage Flight Experiment (USFE) focuses on key technologies necessary to demonstrate the operation of an inherently simple propulsion system with an innovative, state-of-the-art structure. Two key low-cost vehicle elements will be demonstrated - a 10,000 lbf thrust engine and an integrated composite tank structure. The suborbital flight test of the USFE is scheduled for 2001. Preceding the flight tests are two major series of ground tests at NASA Stennis Space Center and a subscale tank development program to identify compatible composite materials and to verify their compatibility over long periods of time. The ground tests include a thrust chamber development test series and an integrated stage test. This paper summarizes the results from the first phase of the thrust chamber development tests and the results to date from the tank material compatibility tests. Engine and tank configurations that meet the goals of the program are described.

Anderson, W. E.↗

NASA NTRS1993

Simulation of sloshing dynamics induced forces and torques actuated on dewar container driven by gravity gradient and jitter accelerations in microgravity

Some experimental spacecraft use superconducting sensors for gyro read-out and so must be maintained at a very low temperature. The boil-off from the cryogenic liquid used to cool the sensors can also be used, as the Gravity Probe B (GP-B) spacecraft does, as propellant to maintain attitude control and drag-free operation of the spacecraft. The cryogenic liquid for such spacecraft is, however, susceptible to both slosh-like motion and non-axisymmetric configurations under the influence of various kinds of gravity jitter and gravity gradient accelerations. Hence, it is important to quantify the magnitude of the liquid-induced perturbations on the spacecraft. We use the example of the GP-B to investigate such perturbations by numerical simulations. For this spacecraft disturbances can be imposed on the liquid by atmospheric drag, spacecraft attitude control maneuvers, and the earth's gravity gradient. More generally, onboard machinery vibrations and crew motion can also create disturbances. Recent studies suggest that high frequency disturbances are relatively unimportant in causing liquid motions in comparison to low frequency ones. The results presented here confirm this conclusion. After an initial calibration period, the GP-B spacecraft rotates in orbit at 0.1 rpm about the tank symmetry axis. For this rotation rate, the equilibrium liquid free surface shape is a 'doughnut' configuration for all residual gravity levels of 10(exp -6) g(sub 0) or less, as shown by experiments and by numerical simulations; furthermore, the superfluid behavior of the 1.8 K liquid helium used in GP-B eliminates temperature gradients and therefore such effects as Marangoni convection do not have to be considered. Classical fluid dynamics theory is used as the basis of the numerical simulations here, since Mason's experiments show that the theory is applicable for cryogenic liquid helium in large containers. To study liquid responses to various disturbances, we investigate and simulate three levels of gravity jitter (10(exp -6), 10(exp -7), and 10(exp -8) g(sub 0)) each at three predominant frequencies (0.1, 1.0, and 10 Hz), combined with a gravity gradient appropriate for the GP-B orbit. Dynamical evolution of sloshing dynamics excited fluid forces and torque fluctuations exerted on the dewar container driven by the combined gravity gradient and jitter accelerations are also investigated and simulated.

Hung, R. J.↗