Showing papers on "Propellant published in 1998"

Liquid-Propellant Rocket Engine Injector Dynamics

Abstract: The dynamic characteristics of liquid-rocket injectors in the presence of intense combustion-chamber and propellant feed-line oscillations are discussed. Liquid-propellant injectors always function in nonsteady e ow environments and are therefore considered as a dynamic component of an engine. In addition to its main function of injecting propellant and preparing a combustible mixture, an injector simultaneously acts as a sensitive element that may generate and modify e ow oscillations because of its intrinsic unsteadiness and interactions with the combustion-chamber and feed-system dynamics. This paper also addresses nonlinear effects of nonstationary processes occurring in injectors. Various mechanisms for driving self-pulsations in both liquid and gas ‐liquid injectors are summarized systematically. A vital problem of bifurcational unsteadiness of injector operation is considered.

Mixing and Spray Formation in Coaxial Jets

Abstract: The e ow regimes and e ne-scale structure of the mixture resulting from the destabilization of a dense, slow jet by a fast light annular jet are discussed. From the primary shear instability between the streams, it is shown that the momentum ratio M = of the annular to the inner stream is the key parameter 2 2 r u /r u 2 2 1 1 on which the inner potential core length and the condition of a recirculation transition depend. The instability analysis shows that the initial wavelength of the disturbances l is proportional to the vorticity thickness d of the fast stream at the nozzle exit l ; d(r1/r 2) 1/2 ; this further provides, by an original scenario, an estimate of the droplets’ size formed by the capillary instability of the sheet’ s rim formed from the initial interface disturbances. Liquid viscosity is not expected to play any signie cant role in practical conditions in liquid rocket propellant engines. These e ndings are put in relation with a selected review of known or yet unexplained results on the subject. HERE is a frequent need to realize a uniform mixture from two initially segregated streams in many practical instances. A signie cant example is the case of liquid rocket propellant engines because the length of the combustion chamber is limited by the ability of the injecting devices to fragment and mix the reactants down to a sufe cient level of homogeneity for evaporation and combustion to completion at the desired distance from the injector outlet. For technological reasons one of the reactants is usually available in the liquid state and the other in the gas phase; for historical reasons the coaxial geometry is commonly used to merge the two streams. 1 The relative e ow rates of the reactants must be adjusted so that the global stoichiometry is at least respected, or such that the gas phase is in excess for all of the liquid to vaporize and burn, this implies that the gas stream is usually much more rapid than the liquid stream at the injector outlet. This is the situation encountered in H 2/O2 engines, where a slow, dense liquid oxygen (LOX) stream in the central jet of the coaxial injector is surrounded by a fast, light, gaseous hydrogen annular stream. At the root of the interpenetration process between the two phases is a strong shear destabilizing the central liquid jet, which further fragments into a more or less uniform spray. This process, which is contrasted with the case of a simple jet issuing in a quiescent environment, 2 is known as airblast atomization. 3

Reservoir volume sensors

Abstract: Reservoir volume in a drug delivery device is sensed by providing a capacitor, the capacitance of which varies with bellows position or, alternatively, with the amount of propellant liquid absorbed in a dielectric material. In one embodiment, a capacitance is provided between a surface of the bellows, which acts as a first capacitor plate, and a conductive surface disposed proximate the bellows, which acts as a second capacitor plate. As the bellows moves from its extended full position to its collapsed empty position, the area of overlap, and therefore the capacitance between the first and second plates varies from a maximum value to a minimum value. In another embodiment, a variable capacitor is provided with an absorbent material. The absorbent material absorbs the liquid phase of the propellant in the pump housing and acts as a dielectric between two stationary conductive plates provided in the housing. The amount of liquid propellant absorbed in the absorbent material varies with the reservoir volume. When the reservoir is in its full, expanded position, more liquid propellant is absorbed in the absorbent material. When the reservoir is in its compressed empty position, more of the propellant exists as vapor within the pump housing and therefore less liquid propellant is absorbed in the absorbent material. The dielectric properties of the capacitor are therefore higher and the capacitance is therefore maximized when the reservoir is in its full, extended position. Conversely, the dielectric properties are lower and the capacitance minimized when the reservoir is in its compact, empty position.

Propellant Losses Because of Particulate Emission in a Pulsed Plasma Thruster

Abstract: : Propellant inefficiency material in particulate form is characterized in a laboratory pulsed plasma thruster (PPT) operating at 1 Hz with a 204 discharge energy (20 W). Exhaust deposits are collected and analyzed using a combination of a scanning electron microscope with energy dispersive x-ray analysis and microscopic imaging. Teflon(trademark) particulates are observed with characteristic dianietens ranging from over 100 micrometers down to less than 1 micrometer.

Propellant technologies : far-reaching benefits for aeronautical and space-vehicle propulsion

Abstract: Rocket propellant and propulsion technology improvements can be used to reduce the development time and operational costs of new space-vehicle programs Advanced propellant technologies can make the space vehicles safer, more operable, and better performing Five technology areas are described: monopropellants, alternative hydrocarbons, gelled hydrogen, metalized gelled propellants, and high-energy density materials The benee ts of these propellants for future vehicles are outlined using mission study results and the technologies are briee y discussed

Propellant Utilization in Hall Thrusters

Abstract: Fuel optimal trajectories for electrically propelled spacecraft require a varying thrust acceleration. To achieve this requirement the electric thruster has to be operated with a variable thrust to power ratio at the maximum available input power. In order to implement this operating mode with a Hall thruster it is necessary to modify both the mass flow rate and the discharge voltage during the flight time. A crucial problem, associated with Hall thruster operation in a variable thrust mode, is the degradation of thruster performance under mass flow rate variations. Experiments with a laboratory Hall thruster have indicated that this degradation is mainly caused by a reduction of the propellant utilization at lower values of the mass flow rate. Modifying the channel geometry, namely, the channel length and channel profile, improvements in the behavior of the propellant utilization at small mass flow rate and consequently an improved thruster performance have been achieved. Results of experimental tests of a Hall thruster at various operating points and geometry are presented.

Investigation of Cryogenic Propellant Flames Using Computerized Tomography of Emission Images

Abstract: Cryogenic propellant combustion is investigated in this paper. It is shown that the mean e ame structure may be obtained by applying computerized tomography principles to oxygen‐ hydrogen (OH) emission images obtained from experiments on a shear coaxial injector. The data correspond to injection conditions typical of those found in rocket motors, but to lower operating pressures of 1, 5, and 10 bar. The transformed emission images yield the mean volumetric OH emission distribution. This quantity may be roughly interpreted as the mean volumetric rate of reaction. The data provide the location of the mean e ame zone and cone rm that stabilization takes place in the immediate vicinity of the injection plane. I. Introduction L IQUID oxygen ‐ gaseous hydrogen rocket engines have been used for a number of years because they yield the high specie c-impulse values needed in space propulsion applications. Cryogenic propellants thus diminish the cost per mass of payload in orbit, but pose specie c storage, handling, and operating problems. Current rocket motor design relies on extensive experience and technological expertise. The detailed processes involved in cyrogenic combustion are, however, not yet fully documented. An improved understanding of the mode of e ame stabilization and of the e ame structure in the near e eld of the injector head would be quite valuable. This information could be used to improve design methodologies and enhance reliability of operation. Such information would be useful for more accurate predictions of heat transfer rates to the engine walls. In this context, knowledge about whether the e ame is stabilized right on the injector lip or at a distance as a lifted e ame is of considerable interest. The stabilization region is specie cally investigated in this paper on the basis of experiments carried out on a cryogenic model scale combustor designated as Mascotte. This facility, operated by ONERA, is dedicated to basic research and technological studies. Data gathered at this facility include planar laser-induced e uorescence (LIF), planar laser light scattering, and emission imaging. Simultaneous recording of light elastic scattering and hydroxyl radical (OH) e uorescence images has allowed identie cation of the e ame stabilization. When the liquid oxygen (LOX) is injected by a central tube and is surrounded by an annulus of high-speed gaseous hydrogen, it is shown 1 that the e ame is established in the outer boundary of the LOX jet, where the hydrogen stream velocity is low. It is also found that the laser-induced OH ‐ e uorescence signal level remains in the same range over the zone visualized, with little change in the signal amplitude as a function of the axial distance. However, the emission images of the excited OH radical appear to yield a different picture of the e ame stabilization region. The emission amplitude is low close to the injector and increases rapidly at a distance. From these specie c features

Multi-chamber inflator

Abstract: A gas generator utilizes two segregated propellant container/combustion chambers, each having a plurality of nonazide propellant grains therein, and each having an igniter for separately igniting the propellant grains located within the respective chambers. A perforated insulating tube disposed between the two chambers precludes flame front and thermal propagation from one chamber to another, thus preventing redeployment of an activated airbag. Alternatively, a gas generator may contain two separate inflator subassemblies insulated to prevent heat conduction from one subassembly to another.

Solid propellant rocket engine

Abstract: aircraft industry. SUBSTANCE: proposed solid-propellant rocket engine has housing accommodating igniter and change in form of solid-propellant grain. Central body of supersonic nozzle is secured on thin-walled reduction tube coupled with movable end face of solid-propellant grain. Guide rod is secured on bottom of housing from side opposite to supersonic nozzle. Supersonic nozzle, spring-loaded relative to housing, is installed in telescopic guide and is rigidly connected with piston arranged in cylinder. Cylinder is secured on housing. Working spaces of cylinder, divided by piston, are connected with engine combustion chamber by channels made in neck of supersonic nozzle with throttling holes and form gas damper with throttling channel made in piston. Thin-walled tube is coupled with movable end face of solid-propellant grain by means of levers whose fixed stop is placed on housing. Short arm of levers rests on solid- propellant grain, and long arm is hinge-connected with thin-walled tube. Damper spaces communicate with supercritical section of supersonic nozzle through outlet throttling holes connected by channel passing in neck of supersonic nozzle. Passage areas of throttling holes of damper rear space exceed those of front space. EFFECT: provision of stable operation of thrust stabilizer with solid-propellant initial temperature open-loop control and combustion chamber, pressure closed-loop control, and rocket acceleration. 3 cl, 3 dwg

Nontoxic orbital maneuvering and reaction control systems for reusable spacecraft

Abstract: Toxic propellants have a high ground operations cost because of the potential hazards that require extensive safety precautions, particularly for reusable spacecraft. Nontoxic propellants for orbital maneuvering and reaction control systems have received periodic attention since the late 1960s as new reusable vehicles and upgrades to existing vehicles are proposed. This paper discusses the spacecraft requirements that drive propellant selections, the viable candidates for nontoxic propellants, and the system concepts and technologies required. Options for nontoxic propellants are also discussed, which are categorized as monopropellants, storable bipropellants, and cryogenic oxygen-based bipropellants. Monopropellants provide inherently simple systems and are most suited to low total impulse systems. Hydrogen peroxide and kerosene is a promising storable bipropellant in terms of its density, specie c impulse, and low toxicity for long-duration spacecraft on-orbit propulsion systems. This combination can be made hypergolic, which renders it a very effective replacement for currently used storable propellants. The hypergolic characteristics of kerosene with additives and high-concentration hydrogen peroxide are presented in detail. Higher-performance liquid oxygen and alcohol or hydrocarbon fuels are advantageous for reusable propulsion systems that emphasize e uid commonality with other spacecraft systems and for human exploration missions where in-situ propellant production is foreseen. The prospects for further research work on all of these propellants are also discussed.

Distributed charge inflator system

Abstract: An inflatable system which uses a fast-burning propellant material (14) distributed within the inflatable component of the system to generate the gas inflating the system. The invention preferably includes a distributed fast-burning igniter material (14), which may be enhanced with additional gas-generating materials to increase the quantities of gas generated, an optional layer, coating, or sheath (17) of supplemental gas-generating material, an environmentally-sealed sheath (12) to protect the enclosed materials from contamination and to improve the burn rates and efficiencies of the propellant and ignition materials, and an electronic squib (11) used to actuate the igniter material upon a signal from the electronic sensor.

Compressed gas propelled aerosol devices

Abstract: A method of reducing the droplet size of a composition sprayed from an aerosol spray device comprising a compressed gas propellant, which method comprises imparting a unipolar charge to the liquid droplets by double layer charging during the spraying of the liquid droplets from the aerosol spray device, the unipolar charge being at a level such that the said droplets have a charge to mass ratio of at least +/- 1 x 10-4 C/kg.

Effect of Propellant Temperature on Efficiency in the Pulsed Plasma Thruster

Abstract: : A pulsed plasma thruster (PPT) benefits from the inherent engineering simplicity-and reduced tankage fraction gained by storing the propellant as a solid. The solid is converted to the gaseous state and accelerated by an electric discharge across the propellant face. Previous research has concluded that as little as 10% of the consumed propellant is converted to plasma and efficiently accelerated. The remaining propellant is consumed in the form of late-time vaporization and particulate emission, creating minimal thrust. Critical to improving the PPT performance is improving the propellant utilization. The present work demonstrates one possible method of increasing the PPT propellant efficiency. By measuring the PPT thrust, propellant consumption, and propellant temperature while varying the power level, duration of the experimental run, and total propellant mass, a correlation is established between decreased propropellant temperature and increased propellant efficiency. The method is demonstrated by performance measurements at 60 W and S W, which show a 25% increase in thrust efficiency, while the propellant temperature decreases from 135 to 42 deg C. Larger increases in the efficiency may be realized on-orbit where operating temperatures are commonly subzero. The dependence of propellant consumption on temperature also creates systematic errors in laboratory measurements with short experimental runs, and orbit analyses where the PPT performance measured at one power level is linearly scaled to the power available on the spacecraft.

Two chamber inflator body

Abstract: A gas generator (10) utilizes two segregated propellant container/combustion chambers (40, 44), each having a plurality of nonazide propellant grains (46, 48) therein, and each having an igniter (58, 60) for separately igniting the propellant grains located within the respective chambers. An insulating disc (20) disposed between the two chambers precludes flame front and thermal propagation from one chamber to another, thus preventing redeployment of an activated airbag.

Solid propellant/water type hybrid gas generator

Abstract: In order to fill an inflatable safety restraint device such as an air bag, a relatively small pyrotechnic charge is used to pressurize a chamber of water or other suitable liquid and force the liquid out of the chamber and through a conduit in a manner wherein it is caused to intimately mix and atomize in a high speed stream of gas resulting from the combustion of the pyrotechnic charge. The heat of the combustion gas is absorbed by the liquid which is then converted into a large volume. The vaporized liquid and combustion gas is used to inflate the air bag with an essentially non-toxic, low temperature, low particulate atmosphere. The resulting combustion gas and liquid mixture may also be used to extinguish a fire.

Influence of Propellant Composition on Drug Delivery from a Pressurized Metered-Dose Inhaler

Abstract: Hydrofluoroalkanes (HFAs) are used to replace chlorofluorocarbons (CFCs) as non-ozone-depleting propellants for pressurized metered-dose inhalers (pMDIs). HFA 134a and HFA 227 are used in combination to precisely manipulate the density and vapor pressure of pMDI formulations. The influence of propellant composition on the dose delivery characteristics of a suspension-based pMDI formulation was investigated. The results showed that as the density of the propellant blends approached the density of the suspended drug particles, the formulation became more physically stable and exhibited the most consistent dose delivery and greatest respirable fraction. The mass median aerodynamic diameter of the aerosolized particles contained in the emitted dose also was decreased by using propellant blends with higher vapor pressures. The performance of a suspension-based pMDI formulation was optimized by varying the propellant composition using HFA 134a and HFA 227.

Hydrogen Peroxide Propulsion for Smaller Satellites

Abstract: As satellite designs shrink, providing maneuvering and control capability falls outside the realm of available propulsion technology. While cold gas has been used on the smallest satellites, hydrogen peroxide propellant is suggested as the next step in performance and cost before hydrazine. Minimal toxicity and a small scale enable bench top propellant preparation and development testing. Progress toward low-cost thrusters and self-pressurizing tank systems is described.

Numerical model of multiphase flows applied to solid propellant combustion in gun systems

Abstract: A variety of gun propulsion systems are currently being investigated and developed by the US Army Research Laboratory (ARL) including those that utilize advanced solid propellant configurations. The chamber of a gun represents a closed system in which propellant is tightly packaged between the breech-end of the chamber, housing an igniter, and the base of the projectile, residing in the gun tube. Upon ignition and burning, the solid propellant in these systems takes on a highly complex structure that includes the dynamics of propellant combustion and various multiphase flow phenomena. Building on a successful history of gun propulsion modeling and simulation, it is a goal of the ARL to develop a next generation interior ballistics model, NGEN, which incorporates general continuum equations along with auxiliary relations into a single code structure that is both modular and readily transportable between computer architectures. Current interest focuses on application of a two-dimensional version of the model to gun charges in which granular propellant is contained in modular units that are loaded into the gun chamber.

N,N'-azobis-nitroazoles and analogs thereof as igniter compounds for use in energetic compositions

Abstract: Novel energetic compounds are provided comprising N,N'-azobis-nitroazoles or analogs thereof. The compounds are useful as igniter materials in a variety of energetic compositions, particularly gas-generating compositions for inflating automobile or aircraft occupant restraint devices. The compounds are also useful in solid rocket propellant compositions, and as primary explosives to be used as detonators, blasting caps, firearms, and the like. Methods for synthesizing the compounds and manufacturing energetic compositions therewith are provided as well.

Preliminary CFD analysis of the vortex hybrid rocket chamber and nozzle flow field

Abstract: A laboratory scale hybrid rocket engine of a unique new design was found to achieve significantly higher fuel regression rates than in classic hybrids operating at similar conditions with similar propellants. Gaseous oxygen was injected through a ring of tangential ports located between the end of the fuel grain and the exit nozzle. A CFD analysis was employed to confirm the overall features of the experimentally observed flow field in the engine. A coaxial, co-swirling vortex flow was shown to develop in the cylindrical combustion chamber. The outer vortex spirals upward along the fuel surface, toward the forward end of the combustion chamber. The flow migrates radially inward as it approaches the head-end of the grain and develops into an inner, downward-spiraling vortex that exits the nozzle. This unusual flow field seemed to augment combustion, and was effective in promoting high regression rates. Regression rates eight times faster than classic hybrids were experimentally demonstrated using both PMM and HTPB fuel grains. A numerical investigation was conducted to determine the influence of design parameters, such as engine contraction ratio and injection velocity on flow-field characteristics. The swirling flow increases the gas velocity, and therefore the convective heat transfer, near the fuel surface. The serpentine flow path also allows for extended travel distance in the engine, which should increase combustion efficiency.

Integral solid booster and hybrid thrust sustaining system and projectile incorporating the same

Abstract: A rocket according to the invention includes a motor having a fuel chamber and an oxidizer tank, a nose portion provided forward of the motor, and a nozzle surrounded by fins provided aft of the motor. The fuel chamber is provided with a relatively central solid propellant, and a hybrid fuel surrounds the solid propellant. The oxidizer tank is filled with a reactant and coupled to the forward end of the fuel chamber. A pathway is provided between the tank and the fuel chamber for the passage of reactant therethrough. At least one of a valve and a barrier is coupled in the pathway to prevent passage of the reactant into the fuel chamber until after the solid propellant is at least partially consumed. After the solid propellant is at least partially consumed, the valve is opened and/or the barrier is removed to permit the passage of reactant into the fuel chamber. The reactant exothermically reacts with the heated hybrid fuel to create a combustive reaction which creates additional and sustained thrust for the rocket. As a result, the rocket has high initial thrust and sustained thrust, enabling the rocket to maintain maximum velocity and reach a relatively higher altitude than possible with either a solid propellant system or a hybrid propellant system alone.

UV, Visible, and Infrared Spectral Emissions in Hybrid Rocket Plumes

Abstract: A survey was made of the spectral emissions from a 2 x 10 inch labscale hybrid rocket motor system The emissions in the Ultraviolet-Visible (300-750 nm), Near Infrared (750-1100 nm), and Mid Infrared (2-16 μm) regions were studied Baseline emissions were found to consist of the sodium and potassium atomic lines, present due to the use of silica phenolic insulators, and the C2, OH, and CH combustion bands Doped fuel studies were performed, using hydroxyl-terminated polybutadiene (HTPB) fuel mixed with metal salts to introduce emitters into the plume Metals studied included manganese, nickel, cobalt, copper, and iron Iron was studied in both the II and III oxidation states Manganese was also used to study the effect of concentration, and indicated that plume emission is quantitative, giving linear output for the range 5 to 40 ppm Overall, the labscale hybrid was found to offer a stable system for plume spectroscopy, whether for direct studies of the hybrid type rocket, or for use in plume simulations of other propulsion systems Introduction The hybrid rocket motor is of interest to the aerospace community for several reasons Two of these, the possible use of hybrids as boosters and their potential usefulness as a plume simulator for other rocket systems, particularly solids, indicate the need for a thorough survey of their spectroscopic emissions For potential use as boosters, which has been discussed as possible alternatives to the current Space Shuttle Solid Rocket Motors (SRM), hybrids must be evaluated as to base heating effects, which are related to total IR emissions The hybrid also appears to be an excellent motor system to use in ground based testing, especially for the general development of optical monitoring and other measurement techniques A properly designed system, built from the ground up for these types of applications, offers an attractive and safe alternative to the use of solid or liquid propellant systems This paper reports the results of a spectral study of such a hybrid system Atomic Emission Spectroscopy An atomic emission system basically consists of a source into which a sample can be introduced, a wavelength selector, and an optical detector The system used in this study is typical of those used for most emission experiments, but the source of emissions, a labscale hybrid motor, is different in several ways from those normally utilized 1 As a base for comparison, the ideal atomic emission source has the following characteristics: 1 Complete atomization of all elements 2 Controllable excitation energy 3 Sufficient excitation energy to excite all elements 4 Inert chemical environment 5 No background 6 Will accept solutions, gases, or solids 7 Tolerant to various solution conditions and solvents 8 Simultaneous multielement analysis 9 Reproducible atomization and excitation conditions 10 Accurate and precise analytical results The hybrid rocket plume does provide a high energy source for the atomization and excitation of the elements involved The plume temperature has been measured to be approximately 2500-2700 degrees C, which is about the same as a hydrogen-air flame The temperatures in the combustion chamber are even higher, on the order of 3000 degrees C, providing more energy for the atomization process The combustion stoichiometry is set to be fuel lean, which normally should provide lower background emissions Rocket plumes in general then do supply an environment which is capable of good atomization and excitation, and can provide these to solids liquids or gases presented to the combustion chamber Also, the motor control system allows precision metering of the oxidant, rigidly setting the operating point of the motor when fired This renders good reproducibility in firing conditions, which, along with fuel dopant seeding homogeneity, results in reproducibility between firings However, it is obvious that a rocket plume falls short on several of these characteristics Since the only method currently available for sample introduction is doping of the fuel grain, this limits us to metals, metal salts, or other solids, which then must not react with the polymer fuel while curing Hybrid motors display a significant amount of particulate matter in the plume, therefore, there is a component of blackbody type background radiation The plume is exposed to the atmosphere, which is not an inert chemical background, but allows additional oxidant to react with exhaust gases Characteristically, hybrid motors have an inherent tendency to pressure oscillations, on the order of 20-60 Hertz This causes fluctuations in plume intensity However, these fluctuations can normally be averaged or integrated out in the detection process, when using the "standard" atomic spectroscopy PMT or array detectors

HAN TEAN (xm-46) mixing gas generator propellant tank pressurizer for launch vehicles and spacecraft

Abstract: A HAN TEAN mixing gas generator pressurization system is proposed for use on pressure-fed rockets. HAN (hydroxyl ammonium nitrate) and TEAN (triethanol ammonium nitrate) are salts dissolved in water to form a single liquid propellant. Combustion of HAN TEAN results in water, carbon dioxide, and nitrogen at high temperatures and pressures. A cryogenic primary working fluid is mixed the HAN TEAN exhaust products to provide warm high-pressure gas to pressurize the rocket's main propellant tanks.

Method and apparatus for pressurized feeding of liquid propellants to a rocket engine

Abstract: A method and system for delivering pressurized propellants to a rocket engine, that has significant advantages over the current state-of-the-art. One of the propellants, the "pressurizing propellant", is at least partially vaporized and the vapor is in pressure communicating relationship with other propellants on board the rocket-propelled vehicle. This vapor pressure pressurizes the propellants to a sufficient degree that they can be charged directly to the rocket engine, or the pressure may be boosted through pumps, if required. Moreover, the pressurized vapor may be used in other applications on board the vehicle, such as orbital adjustment, attitude control, station keeping, and the like. In several embodiments, the propellants are contained in variable volume reservoirs, exemplified by bladders and diaphragms. These variable volume reservoirs are preferably not subjected to tensile stresses when expanded, and are preferably designed for controlled volume reduction, when being drained of propellant. In some embodiments, the propellant reservoirs are contained in a single housing, and in other embodiments the pressurizing propellant may be contained in a separate tank, as long as pressure communicating relationship is maintained between the propellants. Also provided is a method of refueling one vehicle with liquid propellant from another vehicle, using pressure from a more volatile propellant onboard the supply vehicle.