Buoyant anti-slosh system Patent
Submerged fuel tank baffles to prevent sloshing in liquid propellant rocket flight
Evensen, D. A.↗
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A review of film boiling at cryogenic temperatures.
Film boiling occurs in the quenching of metals, the chilling of biological species, the regenerative cooling of rockets, and the cooling down of a cryogenic fuel tank. Occasionally film boiling is also found in a nuclear reactor or in a cryomagnet. Aspects of film boiling involving an unconstrained liquid mass are considered, giving attention to the evaporation time, the Leidenfrost temperature, solid-liquid contacts, the thermal properties of the solid, effects of coating or scale, wettability, the metastable condition, and the velocity effect on drops. Developments discussed with regard to pool boiling are related to vertical surfaces, film boiling from horizontal surfaces, film boiling from a horizontal cylinder, film boiling from a sphere, and film boiling of helium. Processes of film boiling in a channel are also analyzed.
Hsu, Y. Y.↗
Preliminary test results on a compressed multilayer insulation system for a liquid- hydrogen-fueled rocket.
Jacketed multilayer insulation system design, fabrication and compression testing for liquid hydrogen storage tank
MULTILAYER STRUCTURE↗
Optical Cryogenic Tank Level Sensor
Cryogenic fluids play an important role in space transportation. Liquid oxygen and hydrogen are vital fuel components for liquid rocket engines. It is also difficult to accurately measure the liquid level in the cryogenic tanks containing the liquids. The current methods use thermocouple rakes, floats, or sonic meters to measure tank level. Thermocouples have problems examining the boundary between the boiling liquid and the gas inside the tanks. They are also slow to respond to temperature changes. Sonic meters need to be mounted inside the tank, but still above the liquid level. This causes problems for full tanks, or tanks that are being rotated to lie on their side.
Duffell, Amanda↗
Quantity Distance for the Kennedy Space Center Vehicle Assembly Building for Solid Propellant Fueled Launchers
The NASA KSC VAB was built to process Apollo launchers in the 1960's, and later adapted to process Space Shuttles. The VAB has served as a place to assemble solid rocket motors (5RM) and mate them to the vehicle's external fuel tank and Orbiter before rollout to the launch pad. As Space Shuttle is phased out, and new launchers are developed, the VAB may again be adapted to process these new launchers. Current launch vehicle designs call for continued and perhaps increased use of SRM segments; hence, the safe separation distances are in the process of being re-calculated. Cognizant NASA personnel and the solid rocket contractor have revisited the above VAB QD considerations and suggest that it may be revised to allow a greater number of motor segments within the VAB. This revision assumes that an inadvertent ignition of one SRM stack in its High Bay need not cause immediate and complete involvement of boosters that are part of a vehicle in adjacent High Bay. To support this assumption, NASA and contractor personnel proposed a strawman test approach for obtaining subscale data that may be used to develop phenomenological insight and to develop confidence in an analysis model for later use on full-scale situations. A team of subject matter experts in safety and siting of propellants and explosives were assembled to review the subscale test approach and provide options to NASA. Upon deliberations regarding the various options, the team arrived at some preliminary recommendations for NASA.
Stover, Steven↗
Taming Liquid Hydrogen: The Centaur Upper Stage Rocket, 1958-2002
During its maiden voyage in May 1962, a Centaur upper stage rocket, mated to an Atlas booster, exploded 54 seconds after launch, engulfing the rocket in a huge fireball. Investigation revealed that Centaur's light, stainless-steel tank had split open, spilling its liquid-hydrogen fuel down its sides, where the flame of the rocket exhaust immediately ignited it. Coming less than a year after President Kennedy had made landing human beings on the Moon a national priority, the loss of Centaur was regarded as a serious setback for the National Aeronautics and Space Administration (NASA). During the failure investigation, Homer Newell, Director of Space Sciences, ruefully declared: "Taming liquid hydrogen to the point where expensive operational space missions can be committed to it has turned out to be more difficult than anyone supposed at the outset." After this failure, Centaur critics, led by Wernher von Braun, mounted a campaign to cancel the program. In addition to the unknowns associated with liquid hydrogen, he objected to the unusual design of Centaur. Like the Atlas rocket, Centaur depended on pressure to keep its paper-thin, stainless-steel shell from collapsing. It was literally inflated with its propellants like a football or balloon and needed no internal structure to give it added strength and stability. The so-called "pressure-stabilized structure" of Centaur, coupled with the light weight of its high- energy cryogenic propellants, made Centaur lighter and more powerful than upper stages that used conventional fuel. But, the critics argued, it would never become the reliable rocket that the United States needed.
Dawson, Virginia P.↗
Compendium of fracture mechanics problems
Fracture mechanics analysis results are presented from the following structures/components analyzed at Marshall Space Flight Center (MSFC) between 1982 and 1989: space shuttle main engine (SSME), Hubble Space Telescope (HST), external tank attach ring, B-1 stand LOX inner tank, and solid rocket booster (SRB). Results from the SSME high pressure fuel turbopump (HPFTP) second stage blade parametric analysis determine a critical flaw size for a wide variety of stress intensity values. The engine 0212 failure analysis was a time dependent fracture life assessment. Results indicated that the disk ruptured due to an overspeed condition. Results also indicated that very small flaws in the curvic coupling area could propagate and lead to failure under normal operating conditions. It was strongly recommended that a nondestructive evaluation inspection schedule be implemented. The main ring of the HST, scheduled to launch in 1990, was analyzed by safe-life and fail-safe analyses. First safe-life inspection criteria curves for the ring inner and outer skins and the fore and aft channels were derived. Afterwards the skins and channels were determined to be fail-safe by analysis. A conservative safe-life analysis was done on the 270 redesign external tank attach ring. Results from the analysis were used to determine the nondestructive evaluation technique required.
Stallworth, R.↗
Evaluating and Addressing Potential Hazards of Fuel Tanks Surviving Atmospheric Reentry
In order to ensure reentering spacecraft do not pose an undue risk to the Earth's population it is important to design satellites and rocket bodies with end of life considerations in mind. In addition to considering the possible consequences of deorbiting a vehicle, consideration must also be given to the possible risks associated with a vehicle failing to become operational or reach its intended orbit. Based on recovered space debris and numerous reentry survivability analyses, fuel tanks are of particular concern in both of these considerations. Most spacecraft utilize some type of fuel tank as part of their propulsion system. These fuel tanks are most often constructed using stainless steel or titanium and are filled with potentially hazardous substances such as hydrazine and nitrogen tetroxide. For a vehicle which has reached its scheduled end of mission the contents of the tanks are typically depleted. In this scenario the use of stainless steel and titanium results in the tanks posing a risk to people and property do to the high melting point and large heat of ablation of these materials leading to likely survival of the tank during reentry. If a large portion of the fuel is not depleted prior to reentry, there is the added risk of hazardous substance being released when the tank impact the ground. This paper presents a discussion of proactive methods which have been utilized by NASA satellite projects to address the risks associated with fuel tanks reentering the atmosphere. In particular it will address the design of a demiseable fuel tank as well as the evaluation of off the shelf designs which are selected to burst during reentry.
Kelley, Robert L.↗
Integrated launch and emergency vehicle system
A heavy launch vehicle is discussed. The launch vehicle is comprised of an expendable, multi-container, propellant tank that has a plurality of winged booster propulsion modules at one end and a payload supported by adapter structure at the other end. The preferred payload is an entry module that can be adapted for docking to the space station and used as a return vehicle for the space station crew. Additionally, the payload may include communication satellites, supplies, equipment, and/or structural elements for the space station. The winged propulsion modules are released from the expendable propellant tank, in pairs, and they return to Earth in a controlled glide. After a safe landing, at or near the launch site, the modules are prepared for reuse. The rocket engines for each propulsion module are dual-fuel, dual-mode engines and use methane-oxygen and hydrogen-oxygen from the multi-containers of the propellant tank. When the propulsion modules are released from the expendable propellant tank, the rocket engines are moved into the module cargo bay for the return glide flight.
Martin, James A.↗
STS-112 Mission Highlights Resource Tape, Part 1 of 3
This video, Part 1 of 3, shows the activities of the STS-112 crew on flight days 1 - 3. The crew included Commander Jeff Ashby, Pilot Pam Melroy, and Mission Specialists Sandy Magnus, David Wolf, Piers Sellers, and Fyodor Yurchikhin. Flight day 1 begins with an introduction of the astronauts, seen during their pre-flight banquet, and suit-up. The ingress of some of the crew into the Space Shuttle Atlantis is shown. The launch footage includes the view from a camera mounted on the shuttle's external fuel tank, as well as replays. The separation of the shuttle's booster rockets is also shown. On flight day 2 a view of the payload bay and orbiter docking mechanism on Atlantis is shown from a camera on the shuttle's robotic arm. The footage of flight day 3 includes the docking of Atlantis and the International Space Station (ISS), and the exchange of greetings between the two spacecrews. Views of Earth include a pass over the western United States on flight day 2, and a night view of China on flight day 3.
Source record↗
The Space Shuttle
The Space Transportation System (STS) consisting of the Space Shuttle, Spacelab and the Upper Stages is described, together with the Shuttle's projected missions, test programs and costs. The Orbiter, due to be launched late in 1979, is roughly the size of a DC-9, with weight dry empty of about 75,000 kg, and is reusable, as are the two external fuel tanks. The shape and size of the cargo bay (unobstructed cylindrical compartment, 4.6 m in diameter and 18.3 m long), the engine structure (three rocket engines, fueled by liquid oxygen and liquid hydrogen, each capable of developing 1,668,080 N of thrust), the atmospheric condition in the Orbiter's cabin at sea level pressure, the surface insulation material of coated silica-fiber tiles and reinforced carbon-carbon material with a protection capability of up to 1650 C are discussed in detail. The Spacelab, designed by ESA to fit in the Orbiter's cargo bay for scientific research, and the Upper Stages, intended for inserting payloads in high-energy earth orbits, are also analyzed, as are the two-frequency band communication systems and the on board hardware.
Malkin, M. S.↗
Measured Two-Dimensional Damping Effectiveness of Fuel-Sloshing Baffles applied to Ring Baffles in Cylindrical Tanks
Measured two-dimensional damping forces of baffles with various shapes and perforations are presented for fluid conditions representative of those in liquid-fuel rocket vehicles. The effect of amplitude and frequency of fuel sloshing, and surface proximity on baffle damping are shown. Application of the result sin the prediction of damping effectiveness of ring baffles in cylindrical tanks is demonstrated. Finally, some measurements of damping in a free-free cylindrical tank are presented which verify the predictions based on two-dimensional results. Measurements of certain three-dimensional baffles show that they provide greater damping than ring baffles.
Cole, Henry A., Jr.↗
Chapter 12: Materials for Liquid Propulsion Systems
Earth to orbit launch vehicles are propelled by rocket engines and motors, both liquid and solid. This chapter will discuss liquid engines. The heart of a launch vehicle is its engine. The remainder of the vehicle (with the notable exceptions of the payload and guidance system) is an aero structure to support the propellant tanks which provide the fuel and oxidizer to feed the engine or engines. The basic principle behind a rocket engine is straightforward. The engine is a means to convert potential thermochemical energy of one or more propellants into exhaust jet kinetic energy. Fuel and oxidizer are burned in a combustion chamber where they create hot gases under high pressure. These hot gases are allowed to expand through a nozzle. The molecules of hot gas are first constricted by the throat of the nozzle (de-Laval nozzle) which forces them to accelerate; then as the nozzle flares outwards, they expand and further accelerate. It is the mass of the combustion gases times their velocity, reacting against the walls of the combustion chamber and nozzle, which produce thrust according to Newton's third law: for every action there is an equal and opposite reaction. Solid rocket motors are cheaper to manufacture and offer good values for their cost. Liquid propellant engines offer higher performance, that is, they deliver greater thrust per unit weight of propellant burned. They also have a considerably higher thrust to weigh ratio. Since liquid rocket engines can be tested several times before flight, they have the capability to be more reliable, and their ability to shut down once started provides an extra margin of safety. Liquid propellant engines also can be designed with restart capability to provide orbital maneuvering capability. In some instances, liquid engines also can be designed to be reusable. On the solid side, hybrid solid motors also have been developed with the capability to stop and restart. Solid motors are covered in detail in chapter 11. Liquid rocket engine operational factors can be described in terms of extremes: temperatures ranging from that of liquid hydrogen (-423 F) to 6000 F hot gases; enormous thermal shock (7000 F/sec); large temperature differentials between contiguous components; reactive propellants; extreme acoustic environments; high rotational speeds for turbo machinery and extreme power densities. These factors place great demands on materials selection and each must be dealt with while maintaining an engine of the lightest possible weight. This chapter will describe the design considerations for the materials used in the various components of liquid rocket engines and provide examples of usage and experiences in each.
Halchak, John A.↗
Towards Flange-to-Flange Turbopump Simulations for Liquid Rocket Engines
The primary objective of this research is to support the design of liquid rocket systems for the Advanced Space Transportation System. Since the space launch systems in the near future are likely to rely on liquid rocket engines, increasing the efficiency and reliability of the engine components is an important task. One of the major problems in the liquid rocket engine is to understand fluid dynamics of fuel and oxidizer flows from the fuel tank to plume. Understanding the flow through the entire turbopump geometry through numerical simulation will be of significant value toward design. This will help to improve safety of future space missions. One of the milestones of this effort is to develop, apply and demonstrate the capability and accuracy of 3D CFD methods as efficient design analysis tools on high performance computer platforms. The development of the MPI and MLP versions of the INS3D code is currently underway. The serial version of INS3D code is a multidimensional incompressible Navier-Stokes solver based on overset grid technology. INS3D-MPI is based on the explicit massage-passing interface across processors and is primarily suited for distributed memory systems. INS3D-MLP is based on multi-level parallel method and is suitable for distributed-shared memory systems. For the entire turbopump simulations, moving boundary capability and an efficient time-accurate integration methods are build in the flow solver. To handle the geometric complexity and moving boundary problems, overset grid scheme is incorporated with the solver that new connectivity data will be obtained at each time step. The Chimera overlapped grid scheme allows subdomains move relative to each other, and provides a great flexibility when the boundary movement creates large displacements. The performance of the two time integration schemes for time-accurate computations is investigated. For an unsteady flow which requires small physical time step, the pressure projection method was found to be computationally efficient since it does not require any subiterations procedure. It was observed that the artificial compressibility method requires a fast convergence scheme at each physical time step in order to satisfy incompressibility condition. This was obtained by using a GMRES-ILU(0) solver in our computations. When a line-relaxation scheme was used, the time accuracy was degraded and time-accurate computations became very expensive. The current geometry for the LOX boost turbopump has various rotating and stationary components, such as inducer, stators, kicker, hydrolic turbine, where the flow is extremely unsteady. Figure 1 shows the geometry and computed surface pressure of the inducer. The inducer and the hydrolic turbine rotate in different rotational speed.
Kiris, Cetin↗
Supercomputing Aspects for Simulating Incompressible Flow
The primary objective of this research is to support the design of liquid rocket systems for the Advanced Space Transportation System. Since the space launch systems in the near future are likely to rely on liquid rocket engines, increasing the efficiency and reliability of the engine components is an important task. One of the major problems in the liquid rocket engine is to understand fluid dynamics of fuel and oxidizer flows from the fuel tank to plume. Understanding the flow through the entire turbo-pump geometry through numerical simulation will be of significant value toward design. One of the milestones of this effort is to develop, apply and demonstrate the capability and accuracy of 3D CFD methods as efficient design analysis tools on high performance computer platforms. The development of the Message Passage Interface (MPI) and Multi Level Parallel (MLP) versions of the INS3D code is currently underway. The serial version of INS3D code is a multidimensional incompressible Navier-Stokes solver based on overset grid technology, INS3D-MPI is based on the explicit massage-passing interface across processors and is primarily suited for distributed memory systems. INS3D-MLP is based on multi-level parallel method and is suitable for distributed-shared memory systems. For the entire turbo-pump simulations, moving boundary capability and efficient time-accurate integration methods are built in the flow solver, To handle the geometric complexity and moving boundary problems, an overset grid scheme is incorporated with the solver so that new connectivity data will be obtained at each time step. The Chimera overlapped grid scheme allows subdomains move relative to each other, and provides a great flexibility when the boundary movement creates large displacements. Two numerical procedures, one based on artificial compressibility method and the other pressure projection method, are outlined for obtaining time-accurate solutions of the incompressible Navier-Stokes equations. The performance of the two methods is compared by obtaining unsteady solutions for the evolution of twin vortices behind a flat plate. Calculated results are compared with experimental and other numerical results. For an unsteady flow, which requires small physical time step, the pressure projection method was found to be computationally efficient since it does not require any subiteration procedure. It was observed that the artificial compressibility method requires a fast convergence scheme at each physical time step in order to satisfy the incompressibility condition. This was obtained by using a GMRES-ILU(0) solver in present computations. When a line-relaxation scheme was used, the time accuracy was degraded and time-accurate computations became very expensive.
Kwak, Dochan↗
Medium Fidelity Simulation of Oxygen Tank Venting
The item to he cleared is a medium-fidelity software simulation model of a vented cryogenic tank. Such tanks are commonly used to transport cryogenic liquids such as liquid oxygen via truck, and have appeared on liquid-fueled rockets for decades. This simulation model works with the HCC simulation system that was developed by Xerox PARC and NASA Ames Research Center. HCC has been previously cleared for distribution. When used with the HCC software, the model generates simulated readings for the tank pressure and temperature as the simulated cryogenic liquid boils off and is vented. Failures (such as a broken vent valve) can be injected into the simulation to produce readings corresponding to the failure. Release of this simulation will allow researchers to test their software diagnosis systems by attempting to diagnose the simulated failure from the simulated readings. This model does not contain any encryption software nor can it perform any control tasks that might be export controlled.
Sweet, Adam↗
Pad Safety Personnel Launch Support For STS-200
The launch of a space shuttle is a complex and lengthy procedure. There are many places and components to look at and prepare. The components are the orbiter, solid rocket boosters, external tank, and ground equipment. Some of the places are the launch pad, fuel locations, and surrounding structures. Preparations for a launch include equipment checks, system checks, sniff checks for hazardous commodities, and countless walkdowns. Throughout these preparations, pad safety personnel must always be on call. This requires three shifts of multiple people to be ready when needed. Also, the pad safety personnel must be available for the non-launch tasks that are always present for both launch pads
Guarino, Jennifer↗
Lessons Learned with Metallized Gelled Propellants
During testing of metallized gelled propellants in a rocket engine, many changes had to be made to the normal test program for traditional liquid propellants. The lessons learned during the testing and the solutions for many of the new operational conditions posed with gelled fuels will help future programs run more smoothly. The major factors that influenced the success of the testing were propellant settling, piston-cylinder tank operation, control of self pressurization, capture of metal oxide particles, and a gelled-fuel protective layer. In these ongoing rocket combustion experiments at the NASA Lewis Research Center, metallized, gelled liquid propellants are used in a small modular engine that produces 30 to 40 lb of thrust. Traditional liquid RP-1 and gelled RP-1 with 0-, 5-, and 55-wt% loadings of aluminum are used with gaseous oxygen as the oxidizer. The figure compares the thrust chamber efficiencies of different engines.
Source record↗