Hybrid-propellant rocket: Difference between revisions

specific impulse (${\displaystyle I_{sp}}$)

${\displaystyle I_{sp}}$

History

The first work on hybrid rockets was performed in the late 1930s at IG Farben in Germany and concurrently at the California Rocket Society in the United States. Leonid Andrussow, working in Germany, first thought of the idea for a hybrid propellant rocket. O. Lutz, W. Noeggerath, and Andrussow tested a 10 kilonewtons (2,200 lbf) hybrid rocket motor using a coal and gaseous N₂O as the propellants. Oberth also worked on a hybrid rocket motor using LOX as the oxidizer and graphite as the fuel. The high heat of sublimation of carbon prevented these rocket motors from operating efficiently, as it resulted in a negligible burning rate.^[4]

In the 1940s, the California Pacific Rocket Society used LOX in combination with several different fuel types, including wood, wax, and rubber. The most successful of these tests was with the rubber fuel, which is still the dominant fuel in use today. In June 1951, a LOX/rubber rocket was flown to an altitude of 9 kilometres (5.6 mi).^[4]

Two major efforts occurred in the 1950s. One of these efforts was by G. Moore and K. Berman at General Electric. The duo used 90% H₂O₂ (high test peroxide, known as HTP) and polyethylene in a rod and tube grain design. They drew several significant conclusions from their work. The fuel grain had uniform burning. Grain cracks did not affect combustion, like it does with solid rocket motors. No hard starts were observed (a hard start is a pressure spike seen close to the time of ignition, typical of liquid rocket engines). The fuel surface acted as a flame holder, which encouraged stable combustion. The oxidizer could be throttled with one valve, and a high oxidizer to fuel ratio helped simplify combustion. The negative observations were low burning rates and that the thermal instability of peroxide was problematic for safety reasons. Another effort that occurred in the 1950s was development of a reverse hybrid. In a standard hybrid rocket motor, the solid material is the fuel. In a reverse hybrid rocket motor, the oxidizer is solid. William Avery of the Applied Physics Laboratory used jet fuel and ammonium nitrate, selected for their low cost. His O/F ratio was 0.035, which was 200 times smaller than the ratio used by Moore and Berman.^[4]

In 1953 Pacific Rocket Society (est. 1943) was developing the XDF-23, a 4 inches (10 cm) x 72 inches (180 cm) hybrid rocket, designed by Jim Nuding, using LOX and rubber polymer called "Thiokol". They had already tried other fuels in prior iterations including cotton, paraffin wax and wood. The XDF name itself comes from "experimental Douglas fir" from one of the first units.^[5]

In the 1960s, European organizations also began work on hybrid rockets.

AMROC developed the largest hybrid rockets ever created in the late 1980s and early 1990s. The first version of their engine, fired at the Air Force Phillips Laboratory, produced 312,000 newtons (70,000 lbf) of thrust for 70 seconds with a propellant combination of LOX and hydroxyl-terminated polybutadiene (HTPB). The second version of the motor, known as the H-250F, produced more than 1,000,000 newtons (220,000 lbf) of thrust.^[4]

Korey Kline of Environmental Aeroscience Corporation (eAc) first fired a gaseous oxygen and rubber hybrid in 1982 at Lucerne Dry Lake, CA, after discussions on the technology with Bill Wood, formerly with Westinghouse.^[6] The first SpaceShipOne hybrid tests were successfully conducted by Kline and eAc at Mojave, CA.^[7]

In 1994, the

In its simplest form a hybrid rocket consists of a pressure vessel (tank) containing the liquid oxidiser, the combustion chamber containing the solid propellant, and a mechanical device separating the two. When thrust is desired, a suitable ignition source is introduced in the combustion chamber and the valve is opened. The liquid propellant (or gas) flows into the combustion chamber where it is vaporized and then reacted with the solid propellant. Combustion occurs in a boundary layer diffusion flame adjacent to the surface of the solid propellant.

Generally the liquid propellant is the

Combustion

The governing equation for hybrid rocket combustion shows that the regression rate is dependent on the oxidizer mass flux rate, which means the rate that the fuel will burn is proportional to the amount of oxidizer flowing through the port. This differs from a solid rocket motor, in which the regression rate is proportional to the chamber pressure of the motor.^[4]

${\displaystyle {\dot {r}}=a_{o}G_{o}^{n}}$

where ${\displaystyle {\dot {r}}}$ is the regression rate, a_o is the regression rate coefficient (incorporating the grain length), G_o is the oxidizer mass flux rate, and n is the regression rate exponent.^[4]

As the motor burns, the increase in diameter of the fuel port results in an increased fuel mass flow rate. This phenomenon makes the oxidizer to fuel ratio (O/F) shift during the burn. The increased fuel mass flow rate can be compensated for by also increasing the oxidizer mass flow rate. In addition to the O/F varying as a function of time, it also varies based on the position down the fuel grain. The closer the position is to the top of the fuel grain, the higher the O/F ratio. Since the O/F varies down the port, a point called the

• Mechanically simpler - requires only a single liquid propellant resulting in less plumbing, fewer valves, and simpler operations.
• Denser fuel - fuels in the solid phase generally have higher density than those in the liquid phase, reducing overall system volume.
• Metal additives - reactive metals such as aluminium, magnesium, lithium or beryllium can be easily included in the fuel grain increasing specific impulse(${\displaystyle I_{sp}}$), density, or both.
• Combustion instabilities - Hybrid rockets do not typically exhibit high frequency combustion instabilities that plague liquid rockets due to the solid fuel grain breaking up acoustic waves that would otherwise reflect in an open liquid engine combustion chamber.
• Propellant pressurization - One of the most difficult to design portions of a liquid rocket system are the
  turbopumps
  . Turbopump design is complex as it has to precisely and efficiently pump and keep separated two fluids of different properties in precise ratios at very high volumetric flow rates, often cryogenic temperatures, and highly volatile chemicals while combusting those same fluids in order to power itself. Hybrids have far less fluid to move and can often be pressurized by a blow-down system (which would be prohibitively heavy in a liquid rocket) or self-pressurized oxidizers (such as N₂O).
• Cooling - Liquid rockets often depend on one of the propellants, typically the fuel, to cool the combustion chamber and nozzle due to the very high heat fluxes and vulnerability of the metal walls to oxidation and stress cracking. Hybrid rockets have combustion chambers that are lined with the solid propellant which shields it from the product gases. Their nozzles are often graphite or coated in ablative materials similarly to solid rocket motors. The design, construction, and testing of liquid cooling flows is complex, making the system more prone to failure.

Advantages compared with solid rockets

• Higher theoretical ${\displaystyle I_{sp}}$ - Possible due to limits of known solid oxidizers compared to often used liquid oxidizers.
• Less explosion hazard - Propellant grain is more tolerant of processing errors such as cracks since the burn rate is dependent on oxidizer mass flux rate. Propellant grain cannot be ignited by stray electrical charge and is very insensitive to auto-igniting due to heat. Hybrid rocket motors can be transported to the launch site with the oxidizer and fuel stored separately, improving safety.
• Fewer handling and storage issues - Ingredients in solid rockets are often incompatible chemically and thermally. Repeated changes in temperature can cause distortion of the grain. Antioxidants and coatings are used to keep the grain from breaking down or decomposing.
• More controllable - Stop/restart and throttling are all easily incorporated into most designs. Solid rockets rarely can be shut down easily and almost never have throttling or restart capabilities.

Disadvantages of hybrid rockets

Hybrid rockets also exhibit some disadvantages when compared with liquid and solid rockets. These include:

• Oxidizer-to-fuel ratio shift ("O/F shift") - with a constant oxidizer flow-rate, the ratio of fuel production rate to oxidizer flow rate will change as a grain regresses. This leads to off-peak operation from a chemical performance point of view. However, for a well-designed hybrid, O/F shift has a very small impact on performance because ${\displaystyle I_{sp}}$ is insensitive to O/F shift near the peak.
• Low regression-rate (rate at which the solid phase recedes) fuels often drive multi-port fuel grains. Multi-port fuel grains have poor volumetric efficiency and, often, structural deficiencies. High regression rate liquefying fuels developed in the late 1990s offer a potential solution to this problem.^[8]
• Compared with liquid-based propulsion, re-fueling a partially or totally depleted hybrid rocket would present significant challenges, as the solid propellant cannot simply be pumped into a fuel tank. This may or may not be an issue, depending upon how the rocket is planned to be used.

In general, much less development work has been completed with hybrids than liquids or solids and it is likely that some of these disadvantages could be rectified through further investment in research and development.

One problem in designing large hybrid orbital rockets is that turbopumps become necessary to achieve high flow rates and pressurization of the oxidizer. This turbopump must be powered by something. In a traditional liquid-propellant rocket, the turbopump uses the same fuel and oxidizer as the rocket, since they are both liquid and can be fed to the pre-burner. But in a hybrid, the fuel is solid and cannot be fed to a turbopump's engine. Some hybrids use an oxidizer that can also be used as a monopropellant, such as nitromethane or hydrogen peroxide, and so a turbopump can run on it alone. But nitromethane and hydrogen peroxide are significantly less efficient than liquid oxygen, which cannot be used alone to run a turbopump. Another fuel would be needed, requiring its own tank and decreasing rocket performance.

Fuel

Common fuel choices

A reverse hybrid rocket, which is not very common, is one where the engine uses a solid oxidizer and a liquid fuel. Some liquid fuel options are kerosene, hydrazine, and LH₂. Common fuels for a typical hybrid rocket engine include

Hybrid rocket fuel grains can be manufactured via casting techniques, since they are typically a plastic or a rubber. Complex geometries, which are driven by the need for higher fuel mass flow rates, makes casting fuel grains for hybrid rockets expensive and time consuming due in part to equipment costs. On a larger scale, cast grains must be supported by internal webbing, so that large chunks of fuel do not impact or even potentially block the nozzle. Grain defects are also an issue in larger grains. Traditional fuels that are cast are hydroxyl-terminated polybutadiene (HTPB) and paraffin waxes.[9]

Additive manufacturing

Additive manufacturing is currently being used to create grain structures that were not possible to manufacture. Helical ports have been shown to increase fuel regression rates while also increasing volumetric efficiency.^[10] An example of material used for a hybrid rocket fuel is acrylonitrile butadiene styrene (ABS). The printed material is also typically enhanced with additives to improve rocket performance.^[9]

Oxidizer

Common oxidizer choices

Common oxidizers include gaseous or liquid oxygen, nitrous oxide, and hydrogen peroxide. For a reverse hybrid, oxidizers such as frozen oxygen and ammonium perchlorate are used.^{[4]: 405–406}

Proper oxidizer vaporization is important for the rocket to perform efficiently. Improper vaporization can lead to very large regression rate differences at the head end of the motor when compared to the aft end. One method is to use a hot gas generator to heat the oxidizer in a pre-combustion chamber. Another method is to use an oxidizer that can also be used as a monopropellant. A good example is hydrogen peroxide, which can be catalytically decomposed over a silver bed into hot oxygen and steam. A third method is to inject a propellant that is hypergolic with the oxidizer into the flow. Some of the oxidizer will decompose, heating up the rest of the oxidizer in the flow.^{[4]: 406–407}

Hybrid safety

Generally, well designed and carefully constructed hybrids are very safe. The primary hazards associated with hybrids are:

• Pressure vessel failures - Chamber insulation failure may allow hot combustion gases near the chamber walls leading to a "burn-through" in which the vessel ruptures.
• Blow back - For oxidizers that decompose exothermically such as nitrous oxide or
  nitrogen tetroxide
  unless fuel is present in the oxidizer tank.
• Hard starts - An excess of oxidizer in the combustion chamber prior to ignition, particularly for monopropellants such as nitrous oxide, can result in a temporary over-pressure or "spike" at ignition.

Because the fuel in a hybrid does not contain an oxidizer, it will not combust explosively on its own. For this reason, hybrids are classified as having no

Organizations working on hybrids

Commercial companies

In 1998 SpaceDev acquired all of the intellectual property, designs, and test results generated by over 200 hybrid rocket motor firings by the American Rocket Company over its eight-year life. SpaceShipOne, the first private manned spacecraft, was powered by SpaceDev's hybrid rocket motor burning HTPB with nitrous oxide. However, nitrous oxide was the prime substance responsible for the explosion that killed three in the development of the successor of SpaceShipOne at Scaled Composites in 2007.^[11][12] The Virgin Galactic SpaceShipTwo follow-on commercial suborbital spaceplane uses a scaled-up hybrid motor.

SpaceDev was developing the

U.S. Rockets [14] manufactures and deploys hybrids using self-pressurizing nitrous oxide N₂O and HTPB as well as HTP and HTPB. The High Test Hydrogen Peroxide H₂O₂ 86% and Hydroxyl-terminated polybutadiene (HTPB) and aluminum hybrids developed by U.S. Rockets produce a sea level delivered specific impulse (I_sp) of 240, well above the typical 180 of N₂O-HTPB hybrids. In addition to that, they are self-starting, restartable, have considerably lower combustion instability making them suitable for fragile or manned missions such as Bloodhound SSC, SpaceShip Two or SpaceShip Three. The company has successfully tested^[15] and deployed both pressure fed and pump fed versions of the latter HTP-HTPB style. Deliverables to date have ranged from 6 inch to 18 inch diameter, and development units up to 54 inch diameter. The vendor claims scalability to over 5 meters diameter with regression rates approaching solids, according to literature distributed at the November 2013 Defense Advanced Research Projects Agency meeting for XS-1.

Gilmour Space Technologies began testing Hybrid rocket engines in 2015 with both N₂O and HP with HDPE and HDPE wax blends. 2016 testing includes a 5000 Lb HP/PE engine. The company is planning to use hybrids for both sounding and orbital rockets.

Orbital Technologies Corporation (Orbitec) has been involved in some US government-funded research on hybrid rockets including the "Vortex Hybrid" concept.^[16]

Environmental Aeroscience Corporation (eAc)^[17] was incorporated in 1994 to develop hybrid rocket propulsion systems. It was included in the design competition for the SpaceShipOne motor but lost the contract to SpaceDev.

Rocket Crafters Inc. (RCI) builds and tests hybrid rockets at Florida Institute of Technology

Rocket Lab sells hybrid sounding rockets and related technology.

The Reaction Research Society (RRS), although known primarily for their work with liquid rocket propulsion, has a long history of research and development with hybrid rocket propulsion.

Copenhagen Suborbitals, a Danish rocket group, has designed and test-fired several hybrids using N₂O at first and currently LOX. Their fuel is epoxy, paraffin wax, or polyurethane.^[18] The group eventually moved away from hybrids because of thrust instabilities, and now uses a motor similar to that of the V-2 rocket.

Universities

Space Propulsion Group was founded in 1999 by Dr. Arif Karabeyoglu, Prof. Brian Cantwell and others from Stanford University to develop high regression-rate liquefying hybrid rocket fuels. They have successfully fired motors as large as 12.5 in. diameter which produce 13,000 lbf. using the technology and are currently developing a 24 in. diameter, 25,000 lbf. motor to be initially fired in 2010. Stanford University is the institution where liquid-layer combustion theory for hybrid rockets was developed. The SPaSE group at Stanford is currently working with NASA Ames Research Center developing the Peregrine Sounding rocket which will be capable of 100 km altitude.^[19] Engineering challenges include various types of combustion instabilities.^[20] Although the proposed motor was test fired in 2013, the Peregrine program eventually switched to a standard solid rocket for its 2016 debut.

Florida Institute of Technology has successfully tested and evaluated hybrid technologies with their Panther Project. The

Boston University's student-run "Rocket Propulsion Group",^[22] which in the past has launched only solid motor rockets, is attempting to design and build a single-stage hybrid sounding rocket to launch into sub-orbital space by July 2015.^[23]

Brigham Young University (BYU), the University of Utah, and Utah State University launched a student-designed rocket called Unity IV in 1995 which burned the solid fuel hydroxyl-terminated polybutadiene (HTPB) with an oxidizer of gaseous oxygen, and in 2003 launched a larger version which burned HTPB with nitrous oxide.

University of California, Los Angeles's student-run "University Rocket Project" launches hybrid propulsion rockets utilizing Nitrous Oxide as an oxidizer and HTPB as the fuel. They are currently in the development process of their second student-built hybrid rocket engine.

University of Toronto's student-run "University of Toronto Aerospace Team", designs and builds hybrid motor rockets.

In 2016, Pakistan's DHA Suffa University successfully developed ^[24] Raheel-1, hybrid rocket engines in 1 kN class, using paraffin wax and liquid oxygen, thereby becoming the first university run rocket research program in the country.^[25] In India, Birla Institute of Technology, Mesra Space engineering and rocketry department has been working on Hybrid Projects with various fuels and oxidizers.

Pars Rocketry Group from Istanbul Technical University has designed and built the first hybrid rocket engine of Turkey, the rocket engine extensively tested in May 2015.^[26]

A United Kingdom-based team (laffin-gas) is using four N₂O hybrid rockets in a drag-racing style car. Each rocket has an outer diameter of 150mm and is 1.4m long. They use a fuel grain of high-density wound paper soaked in cooking oil. The N₂O supply is provided by Nitrogen-pressurised piston accumulators which provide a higher rate of delivery than N₂O gas alone and also provide damping of any reverse shock.

In Italy one of the leading centers for research in hybrid propellants rockets is CISAS (Center of Studies and Activities for Space) "G. Colombo", University of Padua. The activities cover all stages of the development: from theoretical analysis of the combustion process to numerical simulation using CFD codes, and then by conducting ground tests of small scale and large-scale rockets (up to 20 kN, N₂O-Paraffin wax based motors). One of these engines flew successfully in 2009.

In

There are a number of hybrid rocket motor systems available for amateur/hobbyist use in high-powered model rocketry. These include the popular HyperTek systems[27] and a number of 'Urbanski-Colburn Valved' (U/C) systems such as RATTWorks,^[28] HyperTek,^[29] West Coast Hybrids,^[30] Contrail Rockets,^[31] and Propulsion Polymers.^[32] All of these systems use nitrous oxide as the oxidizer and a plastic fuel (such as Polyvinyl chloride(PVC) or Polypropylene) or a polymer-based fuel such as HTPB. This reduces the cost per flight compared to solid rocket motors, although there is generally more 'GSE' (ground support equipment) required with hybrids.

In popular culture

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An October 26, 2005 episode of the Television show MythBusters entitled "Confederate Rocket" featured a hybrid rocket motor using liquid nitrous oxide and paraffin wax. The myth purported that during the American Civil War, the Confederate Army was able to construct a rocket of this type. The myth was revisited in a later episode entitled Salami Rocket, using hollowed out dry salami as the solid fuel.

In the February 18, 2007 episode of

• Spacecraft propulsion
• Rocketry
• SpaceDev
• Scaled Composites SpaceShipOne
• Space Propulsion Group

References

External links

Wikimedia Commons has media related to Hybrid rocket engines.

• Developing and testing of a 2kN hybrid rocket engine Template:De icon
• Hybrid Rocket Engines, AspireSpace, membership-based organisation (in UK) (particularly see technical papers)
• Portland State Aerospace Society paraffin hybrid links
• "Preliminary Analysis of Hybrid Rockets for Launching Nanosats into LEO" (PDF). Retrieved 2016-01-31. {{cite journal}}: Cite journal requires |journal= (help)
• Hybridrocket, Private page Template:De icon
• WARR, students developing hybrid engines Template:De icon
• University of Padua spin-off