A hybrid-propellant rocket is a rocket with a rocket motor that uses rocket propellants in two different phases: one solid and the other either gas or liquid. The hybrid rocket concept can be traced back to the early 1930s.
Hybrid rockets avoid some of the disadvantages of solid rockets like the dangers of propellant handling, while also avoiding some disadvantages of liquid rockets like their mechanical complexity. Because it is difficult for the fuel and oxidizer to be mixed intimately (being different states of matter), hybrid rockets tend to fail more benignly than liquids or solids. Like liquid rocket engines, hybrid rocket motors can be shut down easily and the thrust is throttleable. The theoretical specific impulse () performance of hybrids is generally higher than solid motors and lower than liquid engines. as high as 400 s has been measured in a hybrid rocket using metalized fuels. Hybrid systems are more complex than solid ones, but they avoid significant hazards of manufacturing, shipping and handling solid rocket motors by storing the oxidizer and the fuel separately.
History The first work on hybrid rockets was performed in the early 1930s at the Soviet Group for the Study of Reactive Motion. Mikhail Klavdievich Tikhonravov was responsible for the first hybrid propelled rocket launch, the GIRD-9, on 17 August 1933, which reached an altitude of . He would later supervise the design of Sputnik I and the Luna programme. In the late 1930s, there was concurrent work done in Germany at IG Farben and in the United States at the California Rocket Society. Leonid Andrussow, working in Germany, theorized hybrid propellant rockets. O. Lutz, W. Noeggerath, and Andrussow tested a hybrid rocket motor using 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. In 1938, the California Rocket Society experimented with a coal and gaseous oxygen hybrid rocket.
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 .
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% high test peroxide (HTP, or H2O2) and polyethylene (PE) 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 the 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.
In 1953 Pacific Rocket Society (est. 1943) was developing the XDF-23, a 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.
In the 1960s, European organizations also began work on hybrid rockets. ONERA, based in France, and Volvo Flygmotor, based in Sweden, developed sounding rockets using hybrid rocket motor technology. The ONERA group focused on a hypergolic rocket motor, using nitric acid and an amine fuel, developing the LEX sounding rocket. The company flew eight rockets: Once in April 1964, three times in June 1965, and four times in 1967. The maximum altitude the flights achieved was over . The Volvo Flygmotor group also used a hypergolic propellant combination. They also used nitric acid for their oxidizer, but used Tagaform (polybutadiene with an aromatic amine) as their fuel. Their flight was in 1969, lofting a payload to .
Meanwhile, in the United States, United Technologies Center (Chemical Systems Division) and Beech Aircraft were working on a supersonic target drone, known as Sandpiper. It used MON-25 (mixed 25% NO, 75% N2O4) as the oxidizer and polymethyl methacrylate (PMM) and Mg for the fuel. The drone flew six times in 1968, for more than 300 seconds and to an altitude greater than . The second iteration of the rocket, known as the HAST, had IRFNA-PB/PMM for its propellants and was throttleable over a 10/1 range. HAST could carry a heavier payload than the Sandpiper. Another iteration, which used the same propellant combination as the HAST, was developed by Chemical Systems Division and Teledyne Aircraft. Development for this program ended in the mid-1980s. Chemical Systems Division also worked on a propellant combination of lithium and FLOx (mixed F2 and O2). This was an efficient hypergolic rocket that was throttleable. The vacuum specific impulse was 380 seconds at 93% combustion efficiency.
American Rocket Company (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 of thrust for 70 seconds with a propellant combination of LOX and hydroxyl-terminated polybutadiene (HTPB) rubber. The second version of the motor, known as the H-250F, produced more than of thrust.
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. The first SpaceShipOne hybrid tests were successfully conducted by Kline and eAc at Mojave, CA. see also Burt Rutan, Scaled Composites, SpaceShipOne.
In 1994, the U.S. Air Force Academy flew a hybrid sounding rocket to an altitude of . The rocket used HTPB and LOX for its propellant, and reached a peak thrust of and had a thrust duration of 16 seconds.
Basic concepts
In its simplest form, a hybrid rocket consists of a pressure vessel (tank) containing the liquid oxidizer, 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 oxidiser (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 oxidizer and the solid propellant is the fuel because solid oxidizers are extremely dangerous and lower performing than liquid oxidizers. Furthermore, using a solid fuel such as Hydroxyl-terminated polybutadiene (HTPB) or paraffin wax allows for the incorporation of high-energy fuel additives such as aluminium, lithium, or metal hydrides.
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.
where is the regression rate, is the regression rate coefficient (incorporating the grain length), is the oxidizer mass flux rate, and is the regression rate exponent.
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 stoichiometric point may exist at some point down the grain.
Properties Hybrid rocket motors exhibit some obvious as well as some subtle advantages over liquid-fuel rockets and solid-fuel rockets. A brief summary of some of these is given below:
Advantages compared with liquid rockets 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 (), 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 – 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