Browse Topic: Liquid propellants

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Fast Reacting Nano-Composite Energetic Materials: Synthesis and Combustion Characterization20AERP12_1112/1/2020
The advent of nanoparticle fuels enables traditionally diffusive controlled reactions to transition towards kinetically dominant reactions, a transition that results in faster reacting formulations that show promise of harnessing the power equivalent to a monomolecular explosive, but packaged as discretely separate fuel and oxidizer composites. Army Research Laboratory, Research Triangle Park, North Carolina Energetic composites are mixtures of solid fuel and oxidizer particles that, when combined, offer higher calorific output than monomolecular explosives. The composites traditionally deliver energy as diffusion limited reactions and, thus, their power available from reaction is much smaller than any explosive. Technology, however, has advanced particle synthesis, and nanoparticles have become more readily available. The advent of nanoparticle fuels enables traditionally diffusive controlled reactions to transition towards kinetically dominant reactions. This transition results in faster reacting formulations that show promise of harnessing the power equivalent to a monomolecular explosive but packaged as discretely separate fuel and oxidizer composites.
Non-Catalytic Ignition System for High-Performance Advanced Monopropellant ThrustersTBMG-287324/1/2018
Anon-catalytic pyrotechnic ignition system was developed for advanced green monopropellant systems operating with hydroxylammonium nitrate (HAN)-based monopropellant AF-M315E. This technology will provide increased performance and new operating regimes for future NASA missions that employ green propulsion systems. The non-catalytic ignition system is a critical component for future high-thrust, in-space rocket engines that utilize AF-M315E, offering longer thruster life and faster startup than catalytic engines. This work demonstrated the pyrotechnic ignition system in proof-of-concept hot-fire tests showing AF-M315E combustion in a workhorse thruster in a laboratory environment.
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.
Wireless Integrated Microelectronic Vacuum Sensor SystemTBMG-1766911/1/2013
NASA Stennis Space Center’s (SSC’s) large rocket engine test facility requires the use of liquid propellants, including the use of cryogenic fluids like liquid hydrogen as fuel, and liquid oxygen as an oxidizer (gases which have been liquefied at very low temperatures). These fluids require special handling, storage, and transfer technology. The biggest problem associated with transferring cryogenic liquids is product loss due to heat transfer. Vacuum jacketed piping is specifically designed to maintain high thermal efficiency so that cryogenic liquids can be transferred with minimal heat transfer.
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.
Liquid Propellant Gas Generation SystemsAIR1343B (Current)6/17/2013
This information report presents a preliminary discussion of liquid propellant gas generation (LPGG) systems. A LPGG system, as used herein, is defined as a system which stores a liquid propellant and, on command, discharges and converts the liquid propellant to a gas. The LPGG system can interface with a gas-to-mechanical energy conversion device to make up an auxiliary power system. Figure 1 shows a block diagram of LPGG system components which include a propellant tank, propellant expulsion system, propellant control and a decomposition (or combustion) chamber. The purpose of this report is to provide general information on the variety of components and system arrangements which can be considered in LPGG design, summarize advantages and disadvantages of various approaches and provide basic sizing methods suitable for initial tradeoff purposes.
A-6C4 Power Sources Committee
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.
Vortex rocket engine reaps the whirlwind13MOMD0301_053/1/2013
Lightweight combustion chamber walls stay cool because hot-burning propellant remains confined to a central cyclone. THE INNER SURFAC ES OF ROCKET ENGINES can be subjected to extremely high heat. Temperatures as high as 3227°C can result from the powerful, thrust-generating exothermal reaction between fuel and oxidizer, which runs hot enough to melt or damage the walls of the combustion chamber. That's why the cores of hot-running rocket engines are made of temperature-resistant materials that often incorporate networks of cooling passages. In these engine designs, vein-like ducts run with low-temperature (sometimes cryogenic) liquid propellants to carry off the heat and thus maintain structural integrity despite the violent conflagration just adjacent.
Ashley, Steven
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
Power Sources for Fluidic ControlsAIR1245B (Current)5/31/2012
This SAE Aerospace Information Report (AIR) presents a review of the types and general characteristics of power sources that may be used to provide the power for gaseous or liquid fluidic control systems. Fluidic definitions, terminology, units and symbols are defined in Reference 2.1.1.
A-6A3 Flight Control and Vehicle Management Systems Cmt
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.
Aerospace Vehicle Cryogenic Duct SystemsARP735A (Current)8/10/2011
This Aerospace Recommended Practice outlines the design, installation, testing and field maintenance criteria for aerospace vehicle cryogenic duct systems. These recommendations are considered currently applicable guides and are subject to revision due to the continuing development within industry.
AC-9 Aircraft Environmental Systems Committee
Catalytic Microtube Rocket IgniterTBMG-107508/1/2011
Devices that generate both high energy and high temperature are required to ignite reliably the propellant mixtures in combustion chambers like those present in rockets and other combustion systems. This catalytic microtube rocket igniter generates these conditions with a small, catalysis-based torch. While traditional spark plug systems can require anywhere from 50 W to multiple kW of power in different applications, this system has demonstrated ignition at less than 25 W. Reactants are fed to the igniter from the same tanks that feed the reactants to the rest of the rocket or combustion system. While this specific igniter was originally designed for liquid methane and liquid oxygen rockets, it can be easily operated with gaseous propellants or modified for hydrogen use in commercial combustion devices.
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.
GAS POWER SERVOS AND REACTION CONTROL SYSTEMSARP739 (Historical)7/16/2008
Various gas systems are classified in a broad sense, component operation is described in moderate detail, pertinent design parameters are discussed, and possible modes for system operation are listed.
A-6B1 Hydraulic Servo Actuation Committee
Liquid Propellant Gas Generation SystemsAIR1343A (Historical)11/6/2007
This information report presents a preliminary discussion of liquid propellant gas generation (LPGG) systems. A LPGG system, as used herein, is defined as a system which stores a liquid propellant and, on command, discharges and converts the liquid propellant to a gas. The LPGG system can interface with a gas-to-mechanical energy conversion device to make up an auxiliary power system. Figure 1 shows a block diagram of LPGG system components which include a propellant tank, propellant expulsion system, propellant control and a decomposition (or combustion) chamber. The purpose of this report is to provide general information on the variety of components and system arrangements which can be considered in LPGG design, summarize advantages and disadvantages of various approaches and provide basic sizing methods suitable for initial tradeoff purposes.
A-6C4 Power Sources Committee
Power Sources for Fluidic ControlsAIR1245A (Historical)1/11/2007
This SAE Aerospace Information Report (AIR) presents a review of the types and general characteristics of power sources that may be used to provide the power for gaseous or liquid fluidic control systems. Fluidic definitions, terminology, units and symbols are defined in Reference 2.1.1.
A-6A3 Flight Control and Vehicle Management Systems Cmt
Next-generation launchers: Only the tried and true need applyAEROJUN06_026/1/2006
Companies find that some elements of the recipe for success have been handed down for generations. When asked how NASA's underdevelopment Crew Launch Vehicle (CLV) can be both “proven” and “cutting edge” at the same time, as it is described in NASA press materials, one official responded: “Cutting edge is in the eye of the beholder.” Cutting edge is fine for other segments of aerospace technology development, but when it comes to space, the watchwords are “experienced,” “legacy,” and “track record,” defining factors for the launcher that NASA expects to take the nation back to the moon and on to Mars.
Rosenberg, Barry
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.
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.
Investigation of Mars In-Situ Propellant Production9724967/1/1997
In-situ production of oxygen and methane for utilization as a return propellant from Mars for both sample-return and manned missions is currently being developed by NASA in cooperation with major aerospace companies. Various technologies are being evaluated using computer modeling and analysis at the system level. An integrated system that processes the carbon dioxide in the Mars atmosphere to produce liquid propellants has been analyzed. The system is based on the Sabatier reaction that utilizes carbon dioxide and hydrogen to produce methane and water. The water is then electrolyzed to produce hydrogen and oxygen. While the hydrogen is recycled, the propellant gases are liquefied and stored for later use. The process model considers the surface conditions on Mars (temperature, pressure, composition), energy usage, and thermal integration effects on the overall system weight and size. Current mission scenarios require a system that will produce 0.7 kg of propellant a day for 500 days. This paper will outline a baseline system for a return propellant production plant, including power requirements and weight estimates, as well as identify key technological improvements required to develop a viable and robust system.
Reddig, Mike, MacKnight, Allen
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
NITROGEN, LIQUID, PROPELLANT GRADEAMS3015 (Historical)5/1/1995
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
AUXILIARY POWER SOURCES FOR AEROSPACE APPLICATIONSAIR744A (Historical)7/1/1992
A-6C4 Power Sources Committee
POWER SOURCES FOR FLUIDIC CONTROLSAIR1245 (Historical)11/1/1991
This AIR concerns itself with the end use of Fluidic (or Flueric) control hardware on aerospace vehicle applications. The fluidic control hardware application is viewed as a system comprised of the following subsystems: Power Source Power Conditioner Fluidic/Flueric Control(s) This AIR identifies potential power sources and relates the design of the fluidic/flueric controls to the nature of both the power source and, as required, the power conditioner. In the unlikely event that the power source yields a fluid which is always at the desired pressure level, temperature range and flow rate capacity and, further, is free of particulate or liquid contaminate, pressure pulsation, etc., no power conditioner is required. Experience has shown that the power conditioner is usually necessary to assure operability and reliability of the total control system. The functions of the fluidic power conditioner are analagous to those of the electrical power supply regulator circuit in an electrical control system or those of the hydraulic servo supply regulator in a hydraulic control system. Examples of (fluidic) power sources are given along with the referencing of government or industry standards (or other documents) which are of interest. AIR 744A "Auxiliary Power Sources for Aerospace Applications" is an excellent, more broadly based document which may be utilized in conjunction with this AIR. Input parameters and output parameters, relative to the power conditioner, are presented. Also, ambients affecting the design of the power conditioner are indicated.
A-6A3 Flight Control and Vehicle Management Systems Cmt
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
INDEX OF STARTING SYSTEM SPECIFICATIONS AND STANDARDSAIR1174 (Historical)7/15/1970
This report lists military and industry specifications and standards which are used in aerospace engine starting systems. Only those hardware standards which have been specifically designed for engine starting systems are listed. Revisions and amendments which are current for these specifications and standards are not listed.
AE-6 Starting Systems and Auxiliary Power Committee
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.
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.
RECOMMENDED MATERIALS AND PRACTICES FOR USE WITH CRYOGENIC PROPELLANTSAIR839A (Historical)4/15/1967
AE-5A Aerospace Fuel, Inerting and Lubrication Sys Committee
RECOMMENDED MATERIALS AND PRACTICES FOR USE WITH CRYOGENIC PROPELLANTSAIR839 (Historical)1/10/1965
AE-5A Aerospace Fuel, Inerting and Lubrication Sys Committee
Propellant Expulsion in Unmanned Spacecraft6407921/1/1964
Bladders, diaphragms, and pistons used for the positive expulsion of earth-storable liquid rocket propellants are discussed in general terms. The history of PL's work on these devices is reviewed as a background to the current programs. A detailed account of the development and use of bladders in Ranger and Mariner spacecraft is presented. The final section describes an advanced development program aimed at providing technology for future spacecraft.
Porter, R. N., Stanford, H. B.
Heat Transfer to Uninsulated Missile Tanks Containing Liquid Hydrogen6303741/1/1963
An approximate analytical method was developed to predict boil-off losses due to external heating of uninsulated missile tanks containing liquid hydrogen. The analysis was applied to the upper-stage tankage of two typical missile configurations which differed only in the type of propulsion utilized; a solid propellant first stage with 90 seconds of action time, and a liquid propellant first stage which burned for 140 seconds. It was predicted that during first-stage flight 15–16% of the liquid hydrogen fuel would be lost due to boil-off. These prohibitive losses could be avoided with a thin layer of light insulation such as corkboard. Therefore, it was concluded that uninsulated missile tanks, containing liquid hydrogen, were not desirable.
Schaechter, Wolfgang
Selection of High Thrust Launch Vehicles6204721/1/1962
NASA and the U.S. Air Force have conducted a number of studies over the past few years to establish the technical feasibility of various large booster concepts and the associated cost, development, and operational characteristics. Many of these studies were performed by various industry agencies; the results then being assembled and interpreted by the cognizant Government agencies. Several of these studies were conducted by North American Aviation, Inc. This paper summarizes some of the significant results of the NASA study of 6-12 million lb thrust launch vehicles (Saturn through Nova Class), contract NAS 8-899, which was conducted under the sponsorship of the Future Projects Office at the Marshall Space Flight Center, Huntsville, Ala. Included are selected study results, a summary of the vehicle types considered and their structural performance, operations, reliability, and economic characteristics. Liquid propellant, solid propellant, and nuclear vehicles are included with both orbital and escape mission capabilities. The material is supplemented to some extent with data from other studies, both in-house and Government sponsored.
Kolyer, R. C.
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.
a study of Self-Contained Starting Systems for Turbojet and Turboprop Engines6000111/1/1960
SUBSTANTIAL POWER is necessary to start the modern jet engine. Thus, starting equipment has become a major concern of air transport operators. This paper discusses the equipment used with self-contained starting systems. The authors discuss and evaluate a variety of self-contained systems: combustor, fuel-air combustion, cartridge, liquid propellant, hydraulic supported by auxiliary power units, and electric supported by APU. Possible future systems are: self-breathing systems, oxygen combustors, and liquid-oxygen-water-fuel combustors. It is emphasized that the choice of a starting system for a particular aircraft will depend on aircraft characteristics and the aircraft's intended use.*
Schmider, Henry R., Ferguson, John H.
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