Browse Topic: Liquid propellants

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Combustion Characteristics of Hydrocarbon Droplets Induced by Photoignition of Aluminum NanoparticlesTBMG-273818/1/2017
In the study of combustion characteristics of liquid rocket fuels, it is customary to either study the combustion of liquid fuel droplets or the combustion of fuel sprays. However, the two are closely related to each other, because in a typical rocket combustion chamber, the burning of droplets, droplet clusters, and fuel sprays occur simultaneously.
Combustion Characteristics of Hydrocarbon Droplets Induced by Photoignition of Aluminum Nanoparticles17AERP08_076/1/2017
Test methodology allows analysis of combustion dynamics for subscale rocket injectors under super critical conditions. Air Force Research Laboratory (AFMC), Edwards Air Force Base, California In the study of combustion characteristics of liquid rocket fuels, it is customary to either study the combustion of liquid fuel droplets or the combustion of fuel sprays. However, the two are closely related to each other, because in a typical rocket combustion chamber, the burning of droplets, droplet clusters, and fuel sprays occur simultaneously. The study of droplet burning is mostly focused either on single suspended droplets or free droplets, usually in a free fall. In either case, the conditions regarding the combustion of a single fuel droplet deviates from what is expected in typical liquid rocket engines (LREs). This not only concerns the chamber pressure and the temperature that is usually much lower in a typical droplet study, but also the droplet ignition process and the related transient effects such as temperature ramp up that is much slower than what takes place in LREs. As a result, reported values of combustion parameters such as the burning rate constant and the ignition delay time obtained under typical experimental conditions may not entirely represent the combustion processes in LREs.
Six-Degree-of-Freedom Control With Only Eight ThrustersTBMG-2346212/1/2015
Typical spacecraft thruster configurations are often unable to provide full six-degree-of-freedom control and may have unwanted interaction between their attitude control and trajectory control functions, have undesirably high instantaneous electrical power demands, and use more thrusters than desirable. These last two potential problems gain increased significance if a spacecraft is required to have especially small size and mass, and have very low cost.
High-Pressure Lightweight ThrustersTBMG-1744510/1/2013
Returning samples of Martian soil and rock to Earth is of great interest to scientists. There were numerous studies to evaluate Mars Sample Return (MSR) mission architectures, technology needs, development plans, and requirements. The largest propulsion risk element of the MSR mission is the Mars Ascent Vehicle (MAV). Along with the baseline solid-propellant vehicle, liquid propellants have been considered. Similar requirements apply to other lander ascent engines and reaction control systems.
Pneumatic Proboscis Heat-Flow ProbeTBMG-163235/1/2013
Heat flow is a fundamental property of a planet, and provides significant constraints on the abundance of radiogenic isotopes, the thermal evolution and differentiation history, and the mechanical properties of the lithosphere. Heat-flow measurements are also essential in achieving at least four of the goals set out by the National Research Council for future lunar exploration. The heat-flow probe therefore directly addresses the goal of the Lunar Geophysical Network, which is to understand the interior structure and composition of the Moon.
Activation and Control of Autoignition in HCCI Engines Using Volumetrically-Distributed Ignition of As-Produced Single-Walled Carbon Nanotubes2012-01-16919/10/2012
The discovery that nanostructured materials exhibit properties different than their bulk materials provided many exciting opportunities with technological applications. One such opportunity is the observed ignition of the single-walled carbon nanotubes (SWCNTs) with an ordinary camera flash. In this paper, light-activated ignition characteristics of the as-produced SWCNTs (50 wt% iron nanoparticle content) with a camera flash are presented. The primary objective of this work is to use nanostructured materials as means for distributed (or volumetric) ignition and improved combustion in propulsion systems. Important examples are homogeneous-charged compression ignition (HCCI) engines, liquid rocket fuel sprays, and enhanced flame stabilization in gas turbine engines. The idea was originally proposed by the author in April 2003 and the first patent filed in July 2004 following a series of initial investigations. Based on these and additional tests, this new ignition method is now considered as a potential enabling technology for volumetric and distributed ignition of liquid fuel sprays or gaseous fuel-air mixtures with the lowest incident power intensity possible. This means remote and spatial ignition within any desired and adjustable region defined by the shape of the light from a pulsed light source. Average intensities in between 10 to 150 W/cm₂ are required for ignition of SWCNTs. This is a factor of 80 less than cases where lasers (pulsed and continuous wave (cw)) are used in coal particles. Results acquired in a premixed gaseous fuel-air mixture in a cylindrical combustion chamber, comparing a spark plug with the light-activated distributed ignition of SWCNTs, confirmed the patented concept and showed a truly on-demand activation of the autoignition process for HCCI engine applications. Faster fuel-air mixture burn rate reaching up to a factor of 3 has been demonstrated for distributed ignition under lean mixture as compared with a conventional spark ignition system.
Chehroudi, Bruce
Oxygen-Methane ThrusterTBMG-132714/1/2012
An oxygen-methane thruster was conceived with integrated igniter/injector capable of nominal operation on either gaseous or liquid propellants. The thruster was designed to develop 100 lbf (≈445 N) thrust at vacuum conditions and use oxygen and methane as propellants. This continued development included refining the design of the thruster to minimize part count and manufacturing difficulties/cost, refining the modeling tools and capabilities that support system design and analysis, demonstrating the performance of the igniter and full thruster assembly with both gaseous and liquid propellants, and acquiring data from this testing in order to verify the design and operational parameters of the thruster.
Technique for Configuring an Actively Cooled Thermal Shield in a Flight SystemTBMG-1136310/1/2011
Broad area cooling shields are a massefficient alternative to conductively cooled thermal radiation shielding. The shield would actively intercept a large portion of incident thermal radiation and transport the heat away using cryogenic helium gas. The design concept consists of a conductive and conformable surface that maximizes heat transfer and formability.
Plasma Igniter for Reliable Ignition of Combustion in Rocket EnginesTBMG-90101/1/2011
A plasma igniter has been developed for initiating combustion in liquid-propellant rocket engines. The device propels a hot, dense plasma jet, consisting of elemental fluorine and fluorine compounds, into the combustion chamber to ignite the cold propellant mixture. The igniter consists of two coaxial, cylindrical electrodes with a cylindrical bar of solid Teflon plastic in the region between them. The outer electrode is a metal (stainless steel) tube; the inner electrode is a metal pin (mild steel, stainless steel, tungsten, or thoriated-tungsten). The Teflon bar fits snugly between the two electrodes and provides electrical insulation between them. The Teflon bar may have either a flat surface, or a concave, conical surface at the open, down-stream end of the igniter (the igniter face). The igniter would be mounted on the combustion chamber of the rocket engine, either on the injector-plate at the upstream side of the engine, or on the sidewalls of the chamber. It also might sit behind a valve that would be opened just prior to ignition, and closed just after, in order to prevent the Teflon from melting due to heating from the combustion chamber.
Self-Adjusting Liquid Injectors for CombustorsTBMG-863310/1/2010
A class of self-adjusting injectors for spraying liquid oxidizers and/or fuels into combustion chambers has been proposed. The proposed injectors were originally intended for use in rocket-engine combustion chambers, but could also be used to improve control over flows of liquid propellants in other combustion chambers.
Self Contained Atmospheric Protective Ensemble (SCAPE) Suits Redesign and Implementation at Kennedy Space Center2005-01-29597/11/2005
The Self Contained Atmospheric Protective Ensemble (SCAPE) suits, worn at the Kennedy Space Center (KSC) have been updated from the original 1970's design. The suits were renamed Propellant Handlers Ensemble (PHE) but are still commonly referred to as SCAPE. Several modifications to the suit were done over the last 20 years to improve the design for operational use. However, anthropometric changes in the user population over time have not been addressed. The following study addressed anthropometric concerns in the current SCAPE population. It was found that all suits had at least one area in which the recommended upper limit was exceeded by technicians. The most common areas to exceed the upper limit were: waist circumference, chest circumference and upper thigh circumference. Forearm circumference posed the least concern unless using long gauntlet glove which cause the twist lock ring to be located at the forearm rather than the wrist.
Mock, Jessica, Miller, Darcy
Precision Stages Enable Analytical Measurement of Jet-Able FluidsTBMG-291614/1/2005
Precision X-Y stages were developed for integration into a new analytical measurement tool for use in the development of digital inks or other fluids to be jetted from ink jet print heads. The Drop Watcher III system provides repeatable and exact measurements of drop formation (e.g., distance and time from the start of the drop ejection), drop size, and flight characteristics of opaque or transparent fluid drops. In addition to X-Y stages, the system consists of a monochrome CCD camera with a zoom lens providing magnification from 0.75X to 4X. Magnification from 2.5X to 10X is achievable when using the high-magnification option, yielding fields of view from 640 µm to 2.56 mm.
Design of a High Effectiveness Micro Exchanger for Mars Applications2000-01-362010/31/2000
NASA is extending human exploration of space beyond the low earth orbit and moon to Mars. To save cost, it has been determined that In Situ Propellant Production (ISPP) is a key enabling technology in Mission to Mars. A cryocooler is needed to liquefy and store oxygen and methane on the Mars surface. In an earlier study by the authors, a single-stage reverse Brayton cycle cryocooler was proposed with neon as the working fluid. The cryocooler operates between 80K and 310K. It was shown that a highly effective recuperative heat exchanger is vital to the overall efficiency of the cryogenic system. To achieve a COP of 0.2 or better, the heat exchanger should have an effectiveness of 0.97 or better while the percentage pressure drop should be less than 3%. In this paper, the design and analysis of a highly effective micro heat exchanger is presented. The heat exchanger is a multi-layer pile of parallel square ducts. The cold and hot fluids flow in a counter flow manner. One-dimensional and three-dimensional numerical models were used to aid in the design. It was determined that in order to achieve an overall pressure drop of 3% or less, the system pressure should be at 4 bars or higher. The need to reduce axial wall conduction to an acceptable level suggests that the heat exchanger should utilize thin wall ducts (20 microns thick) and be constructed with materials with low thermal conductivity (silicon dioxide). The optimal dimensions of the duct are 100 microns for the duct width and 83 mm for the length of the duct.
Zhou, Lei, Kapat, J. S., Chow, L. C., Li, Xiaoyi
Better Densification of Cryogenic Liquid Rocket PropellantsTBMG-3002411/1/1999
An improved system for densifying (by cooling) the liquid hydrogen and liquid oxygen used as propellants for the space shuttle has been proposed. These propellant liquids are cooled to minimize the sizes of tanks needed to store them and to reduce maximum operating pressures. The proposed system would densify these liquids 7 to 8 percent more efficiently than does the present propellant-densification system, and would reduce the absolute vapor pressure from about 15 psi (≈0.1 MPa) to about 1.5 psi (≈10 kPa). Sizing analysis indicates that the combination of the increase in density and the decrease in vapor pressure would reduce the weight of the space shuttle by 10 to 20 percent. Similar reductions in the required volumes and weights of tanks for storing cryogenic liquids could be beneficial in any industry in which the sizes of such tanks are cost factors.
Mars Ascent Propulsion SystemTBMG-299007/1/1999
A report presents an overview of continuing efforts to develop an advanced propulsion system for a spacecraft that would ascend from Mars to bring samples to Earth. The system is required to be smaller and to weigh and cost less, in comparison with a conventionally designed system of equal capability. The development efforts include research on several topics, including the following, which are discussed in the report: (1) warm-gas pressurization subsystems for pushing liquid propellants from supply tanks to engines, (2) lightweight, high-performance rocket engines that burn propellant fluids supplied at temperatures below 0°C, (3) lightweight tanks for propellants and pressurants, and (4) lightweight flow-control components. Finally, the report describes the latest version of the system, which features a two-stage design with pyrotechnic separation.
Hybrid Propulsion System for Returning a Sample From MarsTBMG-322176/1/1998
A paper suggests the development of a hybrid rocket engine and associated equipment for returning a sample of material from Mars at relatively low cost. In a hybrid rocket engine, a solid fuel is burned by use of a liquid or gaseous oxidizer, the flow of which can be throttled to control the engine. Unlike conventional solid rocket propellants, a solid rocket fuel can be made relatively inert in the absence of the oxidizer and therefore presents little hazard of explosion or inadvertent ignition. Unlike conventional (and relatively expensive) liquid rocket propellants, a solid rocket fuel is not corrosive or susceptible to leakage. The solid fuel in the proposed system would be in granular form, packed into the rocket motor. Oxygen or another suitable oxidizer could be transported from Earth together with this solid fuel. Alternatively, oxygen could be generated from CO2 in the Martian atmosphere by use of in-situ resource utilization (ISRU) equipment. Inasmuch as ISRU is not yet a mature technological discipline, some research on ISRU would be necessary to estimate the reduction in cost achieved by not having to carry the oxidizer to Mars.
Nitrogen, Liquid, Propellant GradeAMS3015A (Current)9/1/1996
This specification has been 'CANCELLED' by the Aerospace Materials Division, SAE, as of September 1996.
AMS B Finishes Processes and Fluids Committee
ROCKET IGNITION AND RELATED NOMENCLATUREAIR913 (Current)6/1/1994
This information report provides a short glossary of rocket ignition and related terms.
E-30 Propulsion Ignition Systems Committee
Rocket Propulsion Experimental Facilities at the Phillips Laboratory, Edwards Air Force Base, California9314464/1/1993
The Phillips Laboratory has extensive rocket propulsion facilities at Edwards Air Force Base, California. These facilities range from small scale cells capable of firing liquid rocket engines as small as 0.01 lbf (0.04 Newtons (N)) thrust up to very large liquid/solid test stands designed for firings up to 10 million lbf (44.5 million N) thrust. This paper gives a review of the variety and capabilities of the numerous facilities at Edwards Air Force Base, as well as an overview of current projects. Emphasis is placed on the general capabilities of these facilities; specifics are available upon request to the author. Trends in the evolution of rocket test facilities are discussed.
Beckmann, Joel W., Kalliomaa, Wayne M.
ROCKET ENGINE TEST FACILITIES AT THE U.S. ARMY TEST & EVALUATION COMMAND9314484/1/1993
The Static Test Branch of the Redstone Technical Test Center (RTTC) has developed a unique test capability, in the Army, for static firing hypergolic liquid rocket engines. To accomplish this test requirement, the Army built a dual position test stand for vertical testing of liquid and solid rocket engines. This test stand is supplied with test support facilities both mechanical and electrical in nature. RTTC's facilities and procedures may be used for future testing of Army, DoD, NASA, and private industry's liquid rocket engines.
LILES, WILLIAM M.
Oxygen, Liquid, Propellant GradeAMS3012A (Current)1/1/1990
This specification has been declared "CANCELLED" by the Aerospace Materials Division, SAE as of January 1990.
AMS B Finishes Processes and Fluids Committee
Oxygen, Liquid, Propellant GradeAMS3012 (Historical)4/1/1985
This specification has been declared "CANCELLED" by the Aerospace Materials Division, SAE as of January 1990.
AMS B Finishes Processes and Fluids Committee
Performance of a 10 Millipound Biowaste Resistojet7107692/1/1971
The 10-mlb thrust, concentric-tubes resistojet, which was life tested on hydrogen and ammonia propellants for 8000 hours, is being developed* for use with biowaste propellants. Thrusting performance is presented for H2, H2O, CH4, and CO2 as well as typical biopropellant mixtures CO2-CH4, H2O-CH4 and CO2-H2O. Using Pt-Ir alloy heater tubes, 235 and 575 seconds were obtained for water and hydrogen propellants. Overall total power efficiencies were 66% which includes losses for a pre-evaporator close-coupled for the water tests. Vibration to 1.0 g2/Hz with an overall G rms of 36, shock at 30 g's and acceleration to 8 g's simulating a launch were passed by the thruster.
Phillips, D. G.
Long-Term Storage of Liquid Rocket Propellant Tankage and Components7008002/1/1970
Air Force weapons systems require long-term maintenance-free storage, preferably under uncontrolled environmental conditions. Liquid propulsion system components must be capable of satisfactory operation after years of exposure to highly reactive propellants while retaining the propellant without leakage under severe ambient conditions of temperature and relative humidity. Oxidizer leakage caused by improper component design and severe ambient storage conditions has presented serious operational problems. The Air Force Rocket Propulsion Laboratory (AFRPL) has initiated a program to investigate the storability of liquid system components and tankage under extreme conditions of relative humidity and temperature. A variety of system components and tankage materials are being evaluated for long-term storability with storable liquid rocket fuels and oxidizers. Storage conditions are 85 F temperature and 85% RH for oxidizer systems and +65 to +165 F temperature for fuel systems. The propellants under test are N2O4, C1F5, N2H4, and MHF-5. Tankage materials under test are various alloys of aluminum, steel, and titanium. The results of almost 3 years of testing on a representative number of tankage materials have indicated that leakage of propellant can occur as a result of improper weld joint design, inadequate quality control in fabrication and inadequate acceptance leakage testing. Factors which can contribute to the development of oxidizer leakage are a high ambient relative humidity (30%) and stress-corrosion cracking susceptibility of the tank material in combination with the propellant and trace quantities of foreign compounds/elements in the propellant.
Branigan, John E.
Discrete Element Idealization of Fluids for Vibration Analysis7008472/1/1970
A new type of structural element for discrete element idealization of fluids is investigated. With this structural element and the standard structural elements, a missile and liquid propellant can be idealized as one composite structure consisting of many elements. A theory for taking slosh into account is derived by establishing a slosh stiffness matrix for the fluid elements. This development enables all linear displacement modes of a missile and liquid propellant to be computed. Some of the modes are zero frequency modes, reflecting a true liquid characteristic.
Hunt, D. A.
Space Shuttle Reaction Control System7008022/1/1970
The reaction-control-system requirements for the space shuttle include long life, high reliability, reusability, minimal easy system maintenance, and refurbishment in addition to the usual requirements of reasonable system performance, minimum system weight, and flexibility to operate over a wide range of environmental and operational conditions. The use of gaseous-oxygen/hydrogen propellants potentially provides this capability. Identifiable space-shuttle reaction-control-system requirements are presented, and the two most promising system concepts described and compared. These concepts are the low-chamber-pressure concept, in which the main injection tanks are used as accumulators and heat exchangers to provide the reaction-control engines with the gaseous propellants; and the high-chamber-pressure concept, in which a gas generator/turbopump/heat exchanger is used to provide the propellants to the engines. The overall performances of the two systems are compared and results of sensitivity studies presented. Results of the Manned Spacecraft Center in-house oxygen/hydrogen reaction-control-system-engine technology program are examined, including performance evaluation (combustion efficiency, heating rates, etc.) of three basic injector concepts. Significant factors that influence the engine design relative to application to the space shuttle are discussed.
Pohl, Henry O., Vaughan, Chester A., Kendrick, Darrell
Low Cost Solid Booster Technology6907022/1/1969
This paper discusses the technologies of solid rocket components and the effects of cost consciousness on the state-of-the-art. Motor case technology has developed steels which do not require quench-temper heat treatment and have high flaw tolerance, thus providing major cost optimizing opportunities. Nozzle technology for large solid motors is surprisingly good and accumulating knowledge points to appreciably cheaper nozzles made of prosaic materials being developed in the future. Other areas reviewed in terms of cost appraisal are the case insulation, steering systems, and propellant and processing techniques. In conclusion, possible applications for solid motors, either as prototype systems such as Titan III or intermediate type launch vehicles combining the solid booster first stage with a liquid fueled second stage for low earth orbit or lunar probes are outlined.
Cohen, W., Wasel, R. A.
Flox/Light-Hydrocarbon Propellant Propulsion Research6803622/1/1968
Experimental data on performance and cooling from sea level and altitude firing tests of fluorine/oxygen mixtures with light-hydrocarbon fuels are presented. The majority of the work has been oriented toward pressure-fed rocket applications, with testing at 5000-lb vacuum thrust and 100-psia chamber pressure. Work in progress is at chamber pressures of 250 and 500 psia, being oriented toward pump-fed rocket applications. High delivered specific impulse levels have been demonstrated. One problem, apparently unique to these propellants, is that performance at the optimum mixture ratio is extremely sensitive to incomplete mixing.
Matheson, J. C.
Mars Lander Propulsion6807322/1/1968
Mars landing vehicle descent propulsion is discussed in the light of the constraints of the mission, the environment, and the interfacing functions. The presence of a largely undefined, tenuous atmosphere is shown to produce a great variability of initial and boundary conditions for the propulsion phase of descent. The mission provides numerous subtle constraints of formidable importance to the propulsion implementation. The current requirement for prelaunch, in-situ, thermal sterilization is paramount among these. A significant, and unresolved, mission-related problem is that of the interaction of landing propulsion on the native environment. For a mission which has as its primary objective the investigation of the Mars environment, landing site alteration is a significant concern. The principal intersystem constraint is that of the flight path guidance logic choice. Several options and their propulsion implications are discussed to show the tradeoffs of guidance sophistication and thrust/impulse variability. In this paper, typical Mars propulsive lander mission objectives and constraints are tabulated, the impact of these and other interfaces are reflected in design requirements, and the criteria of design are delineated. A conceptual monopropellant mechanization is presented which satisfies most of these constraints.
Wrobel, J. Richard
Theoretical Performance of Various Chemical Fuels for Auxiliary Power6702052/1/1967
Factors to be considered in selection of fuel for the auxiliary power system are discussed along with the performance of various fuels or propellants. The non-air breathing systems covered in this paper are considered an integrated assembly of four basic subsystems - energy source, energy converter, power conditioning, and heat sink. Advantages and disadvantages of both solid and liquid propellants are reviewed. Because weight and size considerations are of prime importance in auxiliary power units, a theoretical performance in the way of minimum specific propellant consumption is presented. The paper concludes with presentation of a possible problem an APS must undergo, and its solution.
Schultz, R. L.
Navy Reliability Research6706142/1/1967
A study is made of the current research programs in reliability performed in the Navy. Primary emphasis is placed on identifying these programs and requirements for achieving a balanced basic research effort in reliability. Areas of interest include mathematical, engineering, behavioral and management sciences. The two principal functions that a sound program in basic reliability research can fulfill are (1) to increase the emphasis placed upon reliability aspects in relevant scientific research and (2) to provide the fundamental work leading to reliability techniques, methods and knowledge for use across all Naval organizations.
Liebowitz, Harold
Fluid Mechanics of Zero G Propellant Transfer in Spacecraft Propulsion Systems6402201/1/1964
The basic requirement of a transfer system of storable propellants is to provide propellant in sufficient quantity to assure proper engine operation. Various vehicle motions, such as rotational maneuvers, may result in dislocation of propellant from the tank outlet region, thus changing flow rates and mixture ratios and causing impaired performance. In this paper a promising method of propellant relocation, settling by means of an induced gravitational field, is considered.
Gluck, D. F., Gille, J. P.
Vehicle Handling Response to Aerodynamic Inputs6400011/1/1964
The equations of lateral motion response for four wheeled vehicles are developed for external disturbance inputs. Experimental data is obtained through use of a laterally directed hydrogen peroxide rocket motor mounted on a station wagon. The use of a rocket motor provides accurate and flexible control of location and magnitude of the input disturbance. Response data taken from these tests are compared with the responses from a computer model utilizing the disturbance equations. These results are applied to illustrate the effects of wind disturbance on vehicle handling.
Bundorf, R. Thomas, Pollock, Donald E., Hardin, M. C.
The Effects of Cryogenic Fuels on Capacitance Type Mass and Level Measurement6402221/1/1964
Various properties of cryogenic fuels such as ullage gas formation, radial and vertical density gradients, and boiling effects are outlined. The action of these properties on capacitance type mass and level measurement systems is discussed, including the action of ullage gas dielectric, effect of density gradient on probe placement, and the vapor barrier and bubble discontinuities introduced by the boiling effect. Compensation methods and optimum designs for some major cryogenic fluid effects are evaluated. From this evaluation it is concluded that, although present readout systems offer accuracies in the 0.01% area, this accuracy cannot be realized without a complete system design which takes into account the effects of the cryogenic fluid on the system.
Henness, John M., Barnard, Lloyd W.
A Propellant Depletion System for Propulsion Shutdown on the Atlas Space Launch Vehicle6404911/1/1964
In order to establish a propellant depletion system for propulsion shutdown on the Atlas space launch vehicle, design criteria have been closely examined to meet the unique environmental requirements in this particular application. A survey of the different types of liquid level sensing systems in common usage has indicated that the magnetostrictive sensor combines the best qualifications available, and also possesses a high degree of ruggedness and resistance to shock. This paper describes the Atlas space launch vehicle fuel depletion system currently under development and test at General Dynamics/Astronautics. The objective of this program is to produce a system exhibiting the highest possible reliability and performance.
Catlin, Kenneth
Factory-to-Launch Transport and Handling of Future Large Launch Vehicles6303541/1/1963
It is estimated that the space vehicle of tomorrow will stand approximately 445 feet high and weigh over 2-1/2 million pounds when empty. Therefore, the development of suitable techniques to handle and transport these massive space vehicles represents a major challenge to the industries concerned. Special equipment and transportation facilities will have to be utilized in order to deliver the spacecraft and its components from the factory to the launch site. Some of the more important considerations involved are discussed in this paper.
Teague, Beal M.
Engineering Problems of Large Clustered Solid Propellant Boosters6204711/1/1962
A comprehensive investigation of unique design problems of a solid propellant first stage for a typical NOVA launch vehicle is presented. It is found that the major differences between a solid and liquid propellant first stage arise from the higher thrust of solid propulsion units, the cluster structure necessary to combine several motors into a stage, and the vehicle stability and control problem. Analysis of these problem areas shows that satisfactory solutions can be found, using state-of-the-art techniques.
Felix, Bernard R., Wade, Gordon L., Tyson, Mark R.
Orbital Storage of Cryogenic Propellants for Rendezvous Missions6203721/1/1962
This paper discusses analyses which have been conducted on realistic orbital launch booster vehicles utilizing cryogenic propellants. The effect on the design of the cryogenic stages of propellant performance, orbital storage period, and space environment is described, as is the influence of design variations upon the propellant loss rate in orbit of realistic orbital launch boosters. A thermal analysis from boost and transition through orbital conditions is also presented.
Boyles, R. Q., Taton, J. W.
Pressurization Means for Inflatable Structures6203351/1/1962
Results of recent work in the inflation of structures and of flotation and deceleration bags are presented. Solutions to the thermodynamic problems associated with rapid inflation are discussed. Various types of systems are compared with respect to weight, size, and performance. These systems include straight stored gas systems and multiple fluid systems using solid propellants and injectors.
Hebenstreit, L. V., Hadeler, T. T.
Cryogenic Systems in Missiles and Spacecraft6205691/1/1962
Parker, W. F.
A Chemically Fueled Propulsion System for Deep Submergence6103731/1/1961
Howell, H. L.
Some Considerations Affecting the Selection of a Liquid or Solid Propellant Propulsion System6101201/1/1961
Etheridge, F. G.
Current trends in STORABLE PROPELLANTS6000331/1/1960
THE BASIC TREND in storable liquid propellants is directed toward combining high impulse with high-density impulse in low freezing systems. The authors believe that storable liquid propellants with at least 10% higher specific impulses and 20% higher density impulses will be available in the future. Because at present no single oxidizer or fuel exists which combines all the desired attributes of an ideal propellant, the needs of each missile system must be analyzed when choosing a fuel. For example, maximization of specific impulse is increasingly important as the missile range increases. On the other hand, density impulse and wide liquidus range are essential in short-range missiles geared to high mobility prior to launching.*
Silverman, J., Greene, S. A.
some development problems with Large Cryogenic Propellant Systems6000221/1/1960
HEAT TRANSFER causes loading and starting design problems in large missile systems powered by cryogenic propellants. This manifests itself during loading as effective density variation, violent surface conditions, boiloff, and ice formation — problems which may be solved by insulating the tank. During starting it causes overheating and caviation — effects which may be reduced by recirculators and subcooled charge injections. The study described in this paper centers around liquid oxygen and its variations in heat flux rate, which affect liquid density, surface condition, and replenishing requirements. The problem areas are made apparent by consideration of a hypothetical missile system.*
Heald, Daniel A.
SOLID AND LIQUID ROCKETS - A COMPARISON5801641/1/1958
SUTHERLAND, GEORGE S.
THE DESIGN AND TESTING OF SOLID PROPELLANT ROCKET ENGINES5700771/1/1957
VORIS, C. R.
ROCKET POWERPLANTS5201691/1/1952
YOUNG, RAYMOND W.
Fuels For Rocket And Jet Power Plants4702581/1/1947
KLEIN, E. L.
THE ROCKET POWERPLANT 14602281/1/1946
ZUCROW, M. J.
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