Self Pressurized Rockets

A simple propellant feed system for bipropellant liquid rocket engines has been developed. Both propellants are contained within a single tank, and are pressurized by the vapor pressure of one of the propellants.

The pressurization technology has the following characteristics:

· Simple design, with a low parts count and a minimum of active components

· Inexpensive vehicle production, with a minimum of specially fabricated parts or systems

· Very few failure modes, leading to superior reliability

· Stage performance similar to that of conventional pressure fed stages.

· Selectable tank pressures ranging up to approximately 4 MPa (600 psi).

· Reliable propellant acquisition under zero-G conditions

· Pumpless, ventless transfer of propellants under zero-G conditions

· integral provision of gas for cold gas thrusters

The preferred oxidizer for this system is hydrogen peroxide, while the preferred fuel is one of several high vapor pressure liquid hydrocarbons, or mixtures of these hydrocarbons. The vapor from the hydrocarbon pressurizes both the fuel and the oxidizer. The operation of the pressurization system has been modeled analytically, and bench scale hardware validation of system elements has been carried out. The University of British Columbia has been awarded 2 patents on this technology.

BACKGROUND

To avoid the complexity of conventional gas supply systems for pressurizing liquid rocket propellants, there has been some interest in using the vapor pressure of volatile propellants to provide the needed pressure. A number of conventional rocket propellants have sufficiently high critical pressures at room temperature or below to be considered for self pressurization. Cryogenic propellants such as liquid oxygen or liquid methane can be loaded into tanks at their normal boiling point, then warmed until their vapor pressure is high enough for rocket operation. Some potential propellants are high vapor pressure liquids at room temperature. For example, hybrid rockets which utilize a room temperature tank of liquid nitrous oxide and a solid fuel grain have been demonstrated.

Self-pressurized propellants however typically provide poor performance. When raised to a temperature providing sufficient pressure for engine operation, the density of the propellants is considerably lower than normal. Liquid oxygen for example when warmed until it has a vapor pressure of 2 MPa (about 290 psi) has a density of only 0.88, as opposed to a density of 1.14 at its normal boiling point. This increases tank size, which is already a major weight penalty in a pressure fed rocket. Worse, the use of high molecular weight propellants to provide pressurization gas leads to very high burnout masses of tank ullage gas. For most propellants for example, in a stage designed to operate at an initial pressure of 2 MPa, the burnout pressurization gas for self-pressurized systems is about 5 to 8% of the initial propellant mass.

An alternate approach to expelling liquid propellants from a tank is described in a 1983 British patent by Davies, K., Taylor, C.B. and Lewis, J. assigned to British Aerospace. In this concept a tank contains both a non-volatile propellant, and a small amount of a volatile non-propellant liquid (such as ammonia) to provide vapor pressure. The two liquids are separated by a flexible bladder or diaphragm. The volatile non-propellant expels the propellant from the tank by applying pressure to the diaphragm. This system allows a dense, high molecular weight propellant to be expelled by relatively low molecular weight gas. Although this system provides vapor pressure driven expulsion of a non-volatile propellant, it is not suitable for typical rocket applications which require all the propellant to be expelled from the tank in a short time. As the small amount of volatile liquid evaporates, it rapidly cools to the point where its vapor pressure is no longer adequate for rocket operation. This system is therefore restricted to applications such as attitude control systems where propellant is withdrawn in small amounts, and where there is adequate time between withdrawals for the volatile liquid to re-equilibrate in temperature with the rest of the tank.

SELF-PRESSURIZED BIPROPELLANT TECHNOLOGY

Self-pressurized Bipropellant Rockets

One propellant, the "volatile propellant" has a relatively high vapor pressure at the operating temperature of the tank, and partly vaporizes to pressurize the pressure vessel. The volatile propellant is typically a fuel. It is stored in a pressure communicating relationship with a second rocket propellant, which is normally an oxidizer. The latter is typically contained in the same pressure vessel as the fuel, separated by a bladder or diaphragm. A high molecular weight, non-volatile oxidizer thus can be pressurized by a low molecular weight fuel-derived gas. A means is provided to simultaneously draw off the liquid phase of both propellants for use in a pressure fed bipropellant rocket engine. As the propellants are withdrawn, a portion of the volatile propellant evaporates to create high pressure gas to maintain the tank pressure. Pressurant gas can be withdrawn from the ullage space of the tank for provision to cold gas thrusters. Withdrawal of gas has negligible effect on the tank pressure, as withdrawn gas is immediately replenished by vaporization of the liquid phase of the volatile propellant. The figure shows a typical self-pressurized rocket system, using hydrogen peroxide as the oxidizer, and propane as a fuel. A number of other tank configurations are possible, varying in the arrangements for separating the fuel and oxidizer, and in how the fuel is picked up.

Characteristics of Self-pressurized Technology

Self pressurization technology gives particularly simple and low-cost rocket construction as the number of parts in an operational rocket is far fewer than in conventional vehicles. Both propellants are stored in a single tank with exit fittings at the bottom - this eliminates intertank structures and aids packaging. Propellants are non-cryogenic - this allows vehicles to be fueled well in advance of launch, and avoids the problems inherent in cryogenic piping and valving. The low parts count of vehicles strongly aids reliability. Furthermore, the operation of the pressurization system is entirely passive - short of a tank or bladder rupture there is no way in which the system can fail to deliver high pressure propellants to the main engine valves.

Relative to composite case solid rocket motors, self-pressurized rockets have a lower propellant bulk density and require a separate combustion chamber, but have no case insulation and normally are designed to operate at substantially lower pressures (lowering case mass). Thus, self-pressurized rockets can achieve propellant mass fractions similar to those of solid rocket motors, while having substantially higher specific impulse, start-stop capability, built-in pressurized gas for attitude control and an environmentally friendly exhaust.

Zero-G Engine Starts

The propellant with the lower vapor pressure is positively expelled by a bladder or diaphragm system, and thus is available under zero gravity conditions without a propellant settling maneuver. In zero gravity conditions, a propellant settling maneuver is required before main engine ignition to ensure pickup of the higher vapor pressure propellant. This can be done by the vehicle attitude control system. Alternatively, if the lower vapor pressure propellant is a monopropellant (e.g. hydrogen peroxide), the rocket engine may be operated in monopropellant mode first to settle the high vapor pressure propellant prior to bipropellant operation.

Zero G Propellant Transfer

The technology described allows an important secondary benefit - extremely simple and reliable transfer of propellants from one tank to another under zero-G conditions. This has applications for on-orbit refueling of vehicles and scavenging of unused propellants from the upper stage of launchers. Passive and complete transfer of propellants from one tank to another can be done simply by making the donor tank warmer than the receiver tank. The required temperature differential can be achieved by exposing the tanks to the sun while on orbit, and using either differential paint schemes (a black tank and a white tank) or by shading one tank and not the other. If tanks are enclosed within vehicles, one tank would be warmed with an electrical heater. The warmer tank will have the higher pressure. Transfer does not require an external source of pressurized gas to drive propellant from the donor tank, does not require pumps, and does not require venting of gas from the receiving tank. Furthermore, transfer does not require any surface tension propellant acquisition devices in the tanks.

The non-volatile propellant (such as peroxide) is transferred first by opening the valves between the tanks. As the peroxide in the bladder of the donor tank is expelled, the volatile propellant evaporates to keep up the pressure and fill the voided space with gas. As the non-volatile propellant enters the bladder of the receiver tank, the vapor phase of the volatile propellant in that tank condenses under the increased pressure. As the transfer proceeds, the donor tank will get cooler due to volatile propellant evaporation, and the receiver tank will get warmer due to the entry of warm peroxide and to the condensation of volatile propellant. This will lower the pressure differential and slow the transfer as it proceeds. However, as long as the equilibrium temperature of the donor tank is higher than that of the receiver tank, the transfer will go to completion.

Oxidizer (blue) transferred from warm tank (left) to cool tank (right)

Once the non-volatile propellant has been transferred, the volatile propellant (such as a mix of light hydrocarbons) can be moved by opening the appropriate valves. Under zero-G conditions, both gas and liquid may be transferred, depending on the exact behavior of the volatile liquid in zero-G conditions when an outlet to the donor tank is opened. Any liquid residuals in the donor tank will then gradually evaporate from that tank and condense in the receiver tank, due to the difference in equilibrium temperature between the tanks. All that is left in the donor tank after the transfer is a small amount of gas. If a lot of volatile propellant needs to be evaporated from the donor tank, the transfer may take some hours or even days - however, it will go to completion in an entirely passive manner.

Fuel (yellow) transferred from warm tank (left) to cool tank (right)

Applications

Self-pressurized Bipropellant Upper Stages

A number of current launchers utilize relatively expensive and inflexible solid fuel upper stages to provide the final velocity requirements for putting satellites into high orbits, geosynchronous transfer orbits, or interplanetary trajectories. Other vehicles under development, such as the K-1 launcher of Kistler Aerospace by design can only reach low earth orbit, and require a supplementary upper stage for higher energy missions. The technology described here, used with peroxide and a light hydrocarbon fuel, is particular suitable for building low cost upper stages. The principal advantages are that:

· propellants are non-toxic and storable, simplifying ground handling

· propellant are non-cryogenic, allowing multiple engine burns without boiloff losses and eliminating the need for engine temperature pre-conditioning prior to firing

· the peroxide/hydrocarbon combination in a catalyst pack engine has automatic ignition, giving the advantages of conventional hypergolic propellants without their hazards

· if a catalyst pack engine is used with a peroxide lead, no propellant settling maneuver is needed prior to engine burns

· single tank construction gives packaging advantages (no intertank structures or support structures for gas pressurizing bottles)

· the attitude control system can use simple cold gas thrusters supplied from the ullage gas of the main tank, eliminating the need for a separate propellant supply for this system

A Typical Upper Stage

Both fuel and oxidizer are contained within a single pressure vessel, and there are no auxiliary tanks, plumbing, or gas generators. Composite materials such as Kevlar/epoxy or carbon fiber/epoxy are particularly suitable for pressure vessel construction, as they are light for their strength. Because the tank operates near ambient temperature, there is no need for metal construction or special heat-resistant composite materials.

The following table describes a hydrocarbon/peroxide upper stage, operating at an initial pressure of 2 MPa (approx. 290 psi) which decays to a final pressure of approximately 1 MPa at burnout. The hydrocarbon fuel is a mix of ethane and propane giving the desired initial pressure at the stage ambient temperature. The stage is sized to be a direct competitor to a medium sized solid motor such as a Thiokol Star 48. In contrast to the spin stabilized solid stage, the pressure fed liquid stage is 3-axis stabilized by cold gas thrusters.

Mass Characteristics of a Hydrocarbon/Peroxide Upper Stage

	Mass, kg
Tank dry mass (1)	20
Tank burnout pressurization gas (2)	30
Tank fittings and bladder	10
Engine (3)	100
Hardware margin	20
Propellant residuals at 1%	20
Total inert mass	200
Propellant	2,000
Total mass	2,200
Propellant fraction	0.91

(1) Based on performance factor of 300,000 MPa*m³/kg at burst for composite tanks, with safety factor of 1.5

(2) Based on ethane at 15 kg/m³ for 1 MPa and 270 K at burnout.

(3) Based on 50 kN thrust with a specific thrust of 500 N/kg

The stage described above competes directly with standard solid upper stages such as the Star 48 long nozzle. A comparison can be made of the two stages as follows:

Item	Star 48 B	Self-pressurized Bipropellant
Total loaded mass, kg	2141	2200
Inert mass	131	200
Thrust profile	progressive, 60 kN rising to 71	regressive, 50 kN falling to 25
Burn time, seconds	85	180
Expansion ratio	47	40
Specific impulse	292	305
Thrust Vector Control	none	engine gimbal
Stabilization	spin	three axis - cold gas thrusters

Space Station Resupply Missions

For delivery of cargo to a space station, a self-pressurized upper stage would perform the final stages of the rendezvous using cold gas thrusters. This provides a safety advantage over the use of conventional hot gas thrusters, which could cause impingement damage under certain circumstances. After payloads have been delivered to the station, residual propellants in the upper stage would be salvaged for space station use (see above for details of propellant transfer method). Just enough propellant would be left on the upper stage for a de-orbiting burn.

Any unused performance margins from the lower stages of the launcher will appear as unused propellant at the time of rendezvous. In addition, if the delivered payload is less than the nominal capability of the vehicle, the difference will appear as additional unused propellant. The delivered propellant can be stored on orbit at the space station, and used when needed for attitude control (using hydrogen peroxide or cold gas thrusters) or station re-boost (using a small bipropellant engine). Interestingly, over 85% of the mass of the salvageable propellant will be hydrogen peroxide, which can easily be decomposed to water and oxygen. Small turbine auxiliary power units operating on decomposed peroxide could act as a source of electrical power for peaking, emergencies, or base load use when the space station is in shadow. Exhaust gas from the turbine would be run through a condenser coupled to radiator, yielding oxygen and liquid water. Salvaged peroxide is thus not only useful as a propellant, but is a potential source of both electricity and two of the major consumables of the station.

Orbital Launchers

The technology described here is completely scaleable, and could be the basis of a low cost launcher of any desired size. Launch pad facilities can be austere, as the propellants are storable and can be delivered directly to the vehicle from tanker trucks. If peroxide and a light hydrocarbon are used as propellants, the technology is particular suitable for sea launch. In the event of a spill or launch failure, peroxide immediately sinks below a protective layer of water, preventing explosive fuel-oxidizer mixes from forming. Being water miscible, peroxide will rapidly disperse to non-toxic levels as it sinks into deep water. The fuel on the other hand will rapidly evaporate or burn at the water surface, leaving no toxic residue in the water. The high oxidizer to fuel mixture ratio means that a launch vehicle will contain much less flammable fuel than with other propellant combinations.

Because higher chamber pressure can only be achieved at the expense of propellant mass fraction, pressure fed liquid fueled rockets typically have only modest chamber pressures and suffer from considerable loss of specific impulse when operating in the atmosphere. The operational simplicity of self-pressurized rockets however makes them easy to employ in an air launched system where air pressure is low and this problem is not so severe.

TECHNICAL ISSUES

Engine Performance

When using 98% H2O2 as an oxidizer, propane/ethane mixtures used for self-pressurized rockets give a specific impulse approximately 4 to 5 seconds higher than does RP-1 with the same oxidizer. The following graph gives performance values for 98% peroxide and propane/ethane mixes at various chamber pressures, compared with RP-1 (kerosene) burned with peroxide. Chamber pressures for the light hydrocarbon mixes are 75% of the tank pressures for the specified mixes, at 300K. Specific impulse is calculated for full chemical equilibrium to the engine throat, followed by frozen flow expansion in a 40:1 nozzle exhausting to vacuum. It can be seen that light hydrocarbons even at low chamber pressures equal or exceed the performance of RP-1 at higher chamber pressures.

Stage Mass Fraction

In practice, it is estimated that self-pressurized bipropellant rockets will have mass fractions similar to or slightly lower than a competing RP‑1/H2O2 system pressurized with helium, but as discussed above will have slightly higher Isp.

Factors tending to lower the propellant mass fraction of self-pressurized bipropellant rockets

· Self-pressurized bipropellant systems have a lower propellant bulk density than RP-1/peroxide systems, as the hydrocarbon fuel is considerably less dense than kerosene. Typical bulk densities for ethane/propane mixes burned with 98% H2O2 are on the order of 1.0, vs. approximately 1.3 for RP-1/peroxide.

· Self-pressurized bipropellant systems have a pressurizing gas with a molecular weight on the order of 30 to 44 (ethane and propane respectively), whereas conventional pressure fed systems have a pressurizing gas with a molecular weight of 4 (helium).

Factors tending to enhance the propellant mass fraction of self-pressurized bipropellant rockets

· The single tank of the self-pressurized system is structurally efficient, and no intertank brackets or thrust structures are used

· The high molecular weight pressurant of the self-pressurized rocket is stored as a liquid in the same tank it is used, incurring a negligible mass overhead for storage. The high pressure storage bottles, fill valves and regulators of a conventional pressure fed rocket are eliminated, as well as their support structures, valve actuation power supplies and control logic.

Operation and Temperature Effects

Propellant tank pressure at launch depends on the temperature and the precise composition of the fuel mix. For launches in cold weather, the propellants may either be kept warm, or a fuel mix richer in ethane may be used. Overpressurization protection in the event of accidental overheating is provided by a pop-off tank vent.

After ignition, ethane vapor to pressurize the constantly increasing ullage space is provided by evaporation of the ethane (and some propane) from the fuel mix. As the ethane evaporates, the fuel gradually cools and its vapor pressure drops. A further drop in pressure is caused by the selective depletion of the volatile fraction of the remaining liquid fuel. Computer modeling of the process indicates that as the propellants are burned, the final pressure (and thus thrust) is approximately half the initial pressure. This provides automatic throttling of the rocket engine to minimize end acceleration for each stage.

Good pressure control during the flight is provided passively by the characteristics of the system. Throughout the flight, a large surface area of fuel is present. The upper surface of the liquid is always just on the verge of evaporating. Any transient pressure drop will be counteracted by the tendency of the liquid to flash into vapor. Similarly, any transient pressure increase will be counteracted by the tendency of the gas immediately adjacent to the liquid to condense.

The behavior of the volatile propellant during evaporation is expected to mimic what is known to happen in a storage tank of liquefied natural gas when gas is withdrawn. Evaporation happens mainly at the upper surface of the liquid, which is not under hydraulic pressure. The evaporation cools a thin layer of liquid at the upper surface, which then becomes denser and sinks towards the tank bottom. Meanwhile, warmer propellant from underneath rises, and is in turn cooled by evaporation. The tank contents are thus constantly stirred by convection, in an upside-down analog of what happens when a container of water is heated on a hot plate.

When all liquid propellant has been withdrawn from the tank, the tank will still contain a substantial mass of high pressure gas. After the main engine has shut down, this gas may be exhausted through the engine, capturing heat from the still-hot engine and giving a small additional amount of thrust to improve the stage performance.

Propellant Selection

Self-pressurized bipropellant technology is most easily used with rocket propellants maintained at room temperature, where suitable materials for flexible bladders and fuel pickup lines are available. The volatile propellant must have adequate vapor pressure at the operating temperature of the rocket to provide the needed pressurization. Typical lower limits for conventional bipropellant engines are approximately 1 MPa tank pressure. The volatile propellant must have a latent heat of evaporation and heat capacity such that a sufficient fraction of the volatile propellant can be evaporated for use in pressurization without cooling the liquid propellant to the state where its vapor pressure is too low for rocket operation. Finally, the propellants chosen should be non-hypergolic to prevent explosions in the event of bladder failures.

The most suitable volatile propellants are mixtures of low molecular weight hydrocarbons. Mixtures include ethylene or ethane to yield the desired starting vapor pressure, typically 2 MPa. These volatile hydrocarbons are mixed with denser hydrocarbons such as propane, propylene, or methylacetylene (the latter two giving higher Isp than propane). The most favored oxidizer to use with these hydrocarbons is hydrogen peroxide, due to its good performance, high density and lack of nitrogen tetroxide toxicity. Hydrocarbons burned with hydrogen peroxide have a relatively low chamber temperature, facilitating the use of ablative cooling in engines.

Depending on the hydrocarbon mix chosen, rockets operate at tank pressures ranging from 1 MPa or less to 4 MPa (the latter being a pure ethane fuel). Exhaust products are benign and are similar to those of kerosene/LOX engines, but combustion temperature is much cooler. In the event of propellant spills, the hydrocarbon fuel will rapidly and completely evaporate, leaving no environmental damage (in contrast with kerosene which can leave lasting soil residues and which is very toxic in aquatic systems). Hydrogen peroxide spills are easily treated by flooding them with water - dilute peroxide solutions are benign, and in fact 3% hydrogen peroxide is sold in drug stores for cleaning open wounds.

For space use, it will be possible to put a hydrocarbon/peroxide liquid stage into "sleep" mode by radiative cooling. The peroxide will freeze (giving long term stability) and the vapor pressure in the stage will be dropped substantially. When it is desired to use the system, it can be warmed up to melt the peroxide and re-establish a high pressure.

Bladders

Perhaps the most unusual and potentially frightening aspect of the technology described here is the storage of both the fuel and the oxidizer in the same pressure vessel. While this is unusual for liquid rockets, it is standard for solid rockets in which the oxidizer and fuel are intimately mixed and stored in a single casing. There is essentially no pressure differential across the bladder, so even in the event of a small bladder hole there will be little mixing of propellants. Even then, propellants by design aren't mutually miscible and aren't hypergolic - even if they were to contact one another they wouldn't react. Spill simulation tests carried out some decades ago showed for example that it is possible in to float a layer of kerosene on top of a layer of peroxide and even ignite the kerosene without causing an immediate explosion. The lack of any spontaneous reaction between peroxide and hydrocarbons is also witnessed by the common practice of shipping high strength peroxide in polyethylene lined drums.

The propellant bladder is not under tension - the liquid pressure is the same on either side. If correctly designed, the bladder also does not stretch in use, but rather folds as the propellant is removed. The bladder material must be chemically compatible with both the fuel and oxidizer. A starting point for bladder development would be the technology and materials used to make tough fuel bladders for military and industrial storage of diesel fuel and gasoline. Typical fabric weights are 1.5 kg per square meter. If stock fabrics aren't adequately compatible with peroxide, a layer of Teflon or other plastic could be laminated to the bladder surface. Both peroxide and hydrocarbons are routinely stored in polyethylene containers, and polyethylene based bladder materials are thus another possibility, as are bladder materials where the principal polymer is Teflon itself.

Engines

The propellant combination of a light hydrocarbon and hydrogen peroxide is applicable to both reusable and expendable engines. Most peroxide engines to date have used flow-through catalyst packs to induce decomposition of the peroxide. This allows good combustion efficiency, automatic ignition, and restartability. Such engines are particular suitable in small sizes for upper stages. Proven catalysts however (such as silver plated mesh) don't work well with very high strength peroxide, and there is need for development work for catalysts suitable for 90 to 100% peroxide. Assuming the development of a suitable catalyst, light hydrocarbons may have some advantages over kerosene. Light hydrocarbons give a higher specific impulse than kerosene, particularly if they have a positive enthalpy of formation. In addition, combustion efficiency may be superior because of the self-atomizing characteristics of high vapor pressure fuels. Downstream of the pressure drop of the injector, the fuel is superheated with respect to the chamber pressure and will partially flash into vapor as it is injected. This results in a fog of gas and extremely fine liquid droplets which in conjunction with the superheated steam and oxygen from the catalyst pack is likely to have the combustion advantages of gas/gas injection.

For large single-burn engines for use on launchers, alternatives to solid catalyst packs may be desirable. A number of peroxide monopropellant engines have been demonstrated that utilize peroxide plus a small flow of a liquid catalyst such as aqueous potassium permanganate (WW2 German experience as well as recent work by Gary Hudson of HMX). It should be feasible to use a starting slug of such a catalyst in the fuel feed line in place of pyrotechnic ignition.

Because of the wide spread use of catalyst packs with peroxide engines, there has been little research into the use of conventional injectors with this oxidizer. Some work was done by JPL with non-catalytic injectors in the late 1940s and early 1950s. The work was apparently a success and it is possible that details of the results may still be available in JPL archives. One low cost injector design which warrants further investigation is the "coaxial pintle' injector developed by TRW. In this design, one propellant (typically the oxidizer) travels down through a hollow pintle, and is released at low velocity in a radial directions through slots in the periphery of the pintle. The other propellant, typically the fuel, is injected through an annular orifice at the base of the pintle, and travels downwards as a high speed sheet which hits and disperses the low speed oxidizer flow. If a volatile fuel is used in a coaxial pintle engine, the fuel will partially self-vaporize as it passes through the pressure drop of the injector. This may have several benefits, which need to be explored in actual engine tests

· The finely atomized and partly vaporized fuel may have a better combustion efficiency than low volatility fuels such as kerosene. This may allow shorter chamber or chambers with a decreased contraction ratio.

· The fuel stream will accelerate rapidly as it partly bursts into vapor at the flow restriction of the injector. This will give a high speed flow of fluid to violently interact with the oxidizer coming from the pintle, giving good oxidizer dispersion ("perfume spray" mechanism) and again the possibility of improved combustion efficiency.

· A portion of the vapor from the self-vaporization of the injected fuel may form a "radial wind" from the base of the injector which will hit the chamber walls and then form a cool, fuel rich boundary layer which can reduce the need for ablative cooling.

Cold gas thrusters

The operation of the pressurization system naturally generates a large mass of high pressure gas in the ullage space of the tank. This may be tapped for other uses by a fitting at the top of the compartment containing the volatile propellant. For zero-G operations, a liquid blocking device may be used to ensure that only gas is removed from the tank. This could consist of a fine mesh surface tension screen, or a centrifugal separator capable of rejecting liquid.

Withdrawn pressurant may be used in cold gas thrusters for roll control and for stage separation maneuvers (replacing the solid rockets normally used for the latter). Cold gas thrusters utilizing ethane gas have an Isp of approximately 50. In upper stages, ethane may be used in cold gas thrusters for attitude control, propellant settling, minor course corrections, and rendezvous maneuvers. In upper stages, the pressurant available for such uses includes both the tank pressurizing gas at the end of the main engine burn, and unused liquid pressurant from performance margins and mixture ratio biasing (which will evaporate as the depressed tank temperature recovers due to solar heating).

TEST RESULTS

Tank Pressures - Computer Modeling

The relationship between pressure and propellant usage in a self-pressurized system has been modeled independently (using somewhat different approaches) by Dr. Bruce Dunn, and Dr. Phil Hill (University of British Columbia, Mechanical Engineering). Calculations have been performed on a model system in which the volatile propellant is propane, and have for simplicity assumed that there is always complete equilibrium between the liquid and gas phase of the propane. The models by Dr. Dunn and Dr. Hill have provided numerically similar results for the self-pressurized expulsion of propane liquid and oxidizer from a sealed tank.

In a propane powered propellant expulsion system, the initial phase of the blowdown is dominated by the rapid boiling of the bulk propane. The gas thus produced pressurizes the ever-increasing ullage space. Later, when the liquid propane is mostly expelled from the tank, there is little boiling and the system approximates a simple blowdown system in which propellant is expelled by expansion of pre-existing ullage gas.

Typical final tank pressures in a system where the volatile propellant is a pure compound (not a mixture) is 60% of the initial tank pressure. When hydrocarbon mixtures are used, there is an additional vapor pressure drop due to selective depletion of the more volatile of the propellant in the mixture, and the final tank pressure is approximately 50% of the initial tank pressure. In cases where the propellant is used in multiple burns, tank pressure drops are less than these values. Additionally, pressures recover to near initial levels after each burn, as the volatile propellant warms by heat transfer.

Tank Pressures - Test Results

Experiments have been performed to verify the computer blowdown model, using an experimental setup employing propane in modified commercial steel propane tanks, shown in the schematic below. Two tanks rather than a single tank were used, in order to eliminate the need for a bladder in this test apparatus. Propane flow was monitored by the drop in weight of the propane tank, while water (simulated peroxide) flow was monitored by collecting the expelled water in a calibrated receiver. Expelled propane was flared in a burner remote from the test apparatus.

Self-pressurized propellant feed test apparatus

Propane flared from test apparatus during experimental run

The graph below shows the pressure history of the self-pressurized propellant supply system. At the beginning of the run, pressure and propellant flow was less than predicted by the computer model. This appears to reflect superheating of the propane (the presence of liquid propane at a temperature higher than would be predicted by its boiling point at the tank pressure). The delay and the discrepancy between the experimental results and the model could be eliminated by agitating the tank slightly during the run, creating nucleation sites for boiling. In an actual rocket, it is expected that the vibration of the tank would be sufficient to promote boiling and eliminate superheating, and cause the pressure history to follow that predicted by the computer model.

Pressure Vs. Time for Self Pressurized Propellants

Videos of Test Runs

Test Run in Progress (2.3 MB mpg)

Propane flaring, starting with pilot light (2.7 MB mpg)

Flow of simulated oxidizer (2.0 MB mpg)

Technology Status

The technology is covered by two US patents, issued to the University of British Columbia. Potential users of the technology should contact either Dunn Engineering or:

Mr. David Jones
University-Industry Liaison Office
University of British Columbia

Vancouver BC

Canada
(604)-822-8166

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