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Tripropellant rocket
Tripropellant rocket
Tripropellant rocket
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A tripropellant rocket is a rocket that uses three propellants, as opposed to the more common bipropellant rocket or monopropellant rocket designs, which use two or one propellants, respectively. Tripropellant systems can be designed to have high specific impulse and have been investigated for single-stage-to-orbit designs. While tripropellant engines have been tested by Rocketdyne and NPO Energomash, no tripropellant rocket has been flown.

There are two different kinds of tripropellant rockets. One is a rocket engine which mixes three separate streams of propellants, burning all three propellants simultaneously. The other kind of tripropellant rocket is one that uses one oxidizer but two fuels, burning the two fuels in sequence during the flight.

Simultaneous burn

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Simultaneous tripropellant systems often involve the use of a high energy density metal additive, like beryllium or lithium, with existing bipropellant systems. In these motors, the burning of the fuel with the oxidizer provides activation energy needed for a more energetic reaction between the oxidizer and the metal. While theoretical modeling of these systems suggests an advantage over bipropellant motors, several factors limit their practical implementation, including the difficulty of injecting solid metal into the thrust chamber; heat, mass, and momentum transport limitations across phases; and the difficulty of achieving and sustaining combustion of the metal.[1]

In the 1960s, Rocketdyne test-fired an engine using a mixture of liquid lithium, gaseous hydrogen, and liquid fluorine to produce a specific impulse of 542 seconds, likely the highest measured such value for a chemical rocket motor.[2] Despite the high specific impulse, the technical difficulties of the combination and the hazardous nature of the propellants precluded further development.[3]

Sequential burn

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In sequential tripropellant rockets, the fuel is changed during flight, so the motor can combine the high thrust of a dense fuel like kerosene early in flight with the high specific impulse of a lighter fuel like liquid hydrogen (LH2) later in flight. The result is a single engine providing some of the benefits of staging.

For example, injecting a small amount of liquid hydrogen into a kerosene-burning engine can yield significant specific impulse improvements without compromising propellant density. This would have been demonstrated by the RD-701, theoretically capable of a specific impulse of 415 seconds in vacuum (higher than the pure LH2/LOX RS-68), where a pure kerosene engine with a similar expansion ratio would achieve 330–340 seconds.[4]

Although liquid hydrogen delivers the largest specific impulse of the plausible rocket fuels, it also requires huge structures to hold it due to its low density. These structures can weigh a lot, offsetting the light weight of the fuel itself to some degree, and also result in higher drag while in the atmosphere. While kerosene has lower specific impulse, its higher density results in smaller structures, which reduces stage mass, and furthermore reduces losses to atmospheric drag. In addition, kerosene-based engines generally provide higher thrust, which is important for takeoff, reducing gravity drag. So in general terms there is a "sweet spot" in altitude where one type of fuel becomes more practical than the other.

Traditional rocket designs use this sweet spot to their advantage via staging. For instance the Saturn Vs used a lower stage powered by RP-1 (kerosene) and upper stages powered by LH2. Some of the early Space Shuttle design efforts used similar designs, with one stage using kerosene into the upper atmosphere, where an LH2 powered upper stage would light and go on from there. The later Shuttle design is somewhat similar, although it used solid rockets for its lower stages.

SSTO rockets could simply carry two sets of engines, but this would mean the spacecraft would be carrying one or the other set "turned off" for most of the flight. With light enough engines this might be reasonable, but an SSTO design requires a very high mass fraction and so has razor-thin margins for extra weight.

At liftoff the engine typically burns both fuels, gradually changing the mixture over altitude in order to keep the exhaust plume "tuned" (a strategy similar in concept to the plug nozzle but using a normal bell), eventually switching entirely to LH2 once the kerosene is burned off. At that point the engine is largely a straight LH2/LOX engine, with an extra fuel pump hanging onto it.

The concept was first explored in the US by Robert Salkeld, who published the first study on the concept in Mixed-Mode Propulsion for the Space Shuttle, Astronautics & Aeronautics, which was published in August 1971. He studied a number of designs using such engines, both ground-based and a number that were air-launched from large jet aircraft. He concluded that tripropellant engines would produce gains of over 100% (essentially more than double) in payload fraction, reductions of over 65% in propellant volume and better than 20% in dry weight. A second design series studied the replacement of the Shuttle's SRBs with tripropellant based boosters, in which case the engine almost halved the overall weight of the designs. His last full study was on the Orbital Rocket Airplane which used both tripropellant and (in some versions) a plug nozzle, resulting in a spaceship only slightly larger than a Lockheed SR-71, able to operate from traditional runways.[5]

Tripropellant engines were built in Russia. Kosberg and Glushko developed a number of experimental engines in 1988 for a SSTO spaceplane called MAKS, but both the engines and MAKS were cancelled in 1991 due to a lack of funding. However, Glushko's RD-701 was never built, and only a smaller-scale test stand version was tested during development.[6] However, Energomash feels that the problems are entirely solvable and that the design does represent one way to reduce launch costs by about 10 times.[4]

References

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Tripropellant rocket

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from Grokipedia
A tripropellant rocket is a system that utilizes three distinct propellants—typically two and a single oxidizer—to generate , offering potentially higher and efficiency than conventional bipropellant engines by optimizing through simultaneous or burning. These engines fall into two primary categories: one involving the continuous mixing of three liquid propellants or the addition of a solid metal to enhance release, and another employing mode-switching during flight, such as transitioning from a dense for liftoff to for operations. Development of tripropellant concepts dates to the early , with the U.S. and sponsoring investigations into combinations like // and // for their promise of gains of 25 to 69 seconds over bipropellant baselines. A notable example is the Rocketdyne tripropellant engine tested from to 1969 at the , which used , , and to achieve a vacuum of 509 seconds at 95% . The program faced challenges including high heat fluxes exceeding 10 Btu/in²/sec and material erosion, and was ultimately discontinued due to the extreme toxicity and handling difficulties of . In the Soviet Union, the proposed RD-701 engine for the MAKS air-launched spaceplane featured a staged combustion cycle with liquid oxygen, RP-1 kerosene, and liquid hydrogen, delivering up to 3,780 kN of thrust and a vacuum specific impulse of 450–460 seconds across modes, but it remained a design concept without full-scale testing or flight. Subsequent U.S. studies in the 1990s confirmed tripropellant's advantages for single-stage-to-orbit vehicles, including reduced dry mass and increased payload through improved propellant density and impulse, as seen in LOX/LH2/hydrocarbon configurations. A 2015 patent outlined a throttleable tripropellant design with a dual-sleeve pintle injector for flexible propellant combinations, such as kerosene or methane with liquid oxygen, enabling in-flight switching and reusability while addressing cooling issues in oxidizer-rich shutdowns. More recently, in December 2023, Japan's Innovative Space Carrier Inc. achieved the nation's first static fire test of a tripropellant rocket using hydrogen, methane, and oxygen, sustaining combustion for 10 seconds at the Hokkaido Spaceport to validate efficiency for low-cost, reusable small launch vehicles under a government-backed program. In 2024, the company advanced development through partnerships for additive manufacturing of propellant tanks and rocket engines, and presented progress on tripropellant systems for reusable single-stage-to-orbit vehicles at the International Astronautical Congress. Despite these advances, tripropellant systems have not entered operational use due to added complexity, safety risks from corrosive or toxic propellants, and the maturity of high-performance bipropellant alternatives like LOX/methane engines.

Fundamentals

Definition and Classification

A tripropellant rocket is a type of chemical propulsion system that employs three distinct propellants, typically consisting of one fuel paired with two oxidizers or two fuels paired with one oxidizer, to generate thrust through combustion. These propellants are either combusted simultaneously or utilized in sequence, aiming to deliver superior performance metrics compared to traditional systems by optimizing energy release and exhaust properties. Common configurations include a liquid fuel like hydrogen with a solid metal fuel such as beryllium or lithium and a liquid oxidizer like oxygen or fluorine. Tripropellant rockets are classified primarily into two modes based on utilization: simultaneous and sequential. In the simultaneous mode, all three propellants are mixed and burned together within a single , enabling complex reactions that can enhance overall through integrated processes. The sequential mode, by contrast, involves switching between bipropellant combinations derived from the three propellants during flight, such as initially burning a denser fuel-oxidizer pair for high-thrust ascent and transitioning to a higher- pair for operations. This classification allows for mission-specific adaptations, though sequential modes introduce additional engineering complexities in management. The basic components of a tripropellant rocket include separate storage tanks for each , dedicated feed systems with pumps or pressurization mechanisms to deliver the fluids, and a designed to accommodate multiple injection points and reaction zones. Advanced injectors, such as dual-sleeve designs, facilitate precise control over mixing ratios and flow rates to prevent or incomplete . These elements must be robustly engineered to handle the chemical incompatibilities and thermal stresses arising from diverse properties. In comparison to monopropellant systems, which rely on the catalytic decomposition of a single to produce without a separate oxidizer, tripropellant rockets offer far greater and controllability through multi-component reactions. Bipropellant rockets, utilizing one fuel and one oxidizer burned in stoichiometric proportions, provide reliable performance but are limited in flexibility across varying mission phases; tripropellants extend this by enabling optimized (Isp) tailoring, potentially increasing Isp by 13 to 69 seconds over bipropellant baselines like hydrogen-oxygen. This adaptability supports enhanced payload capacities in single-stage-to-orbit applications, albeit at the expense of increased system complexity.

Thermodynamic Principles

Tripropellant rockets employ three distinct propellants to optimize performance metrics such as and density, surpassing the limitations of bipropellant systems by tailoring characteristics to mission phases. Common combinations include and as fuels with as the oxidizer (Li/H₂/F₂), selected for their high energy release and ability to achieve elevated exhaust velocities through low molecular weight products. Another prevalent triad is (LOX) as the oxidizer with (RP-1) and as fuels (LOX/RP-1/H₂), chosen to leverage kerosene's high for sea-level while transitioning to hydrogen's superior . These selections balance propellant for structural and (Isp) for overall velocity increment, with providing substantial volumetric energy in the Li/H₂/F₂ system despite its moderate of approximately 0.534 g/cm³. The specific impulse, a key measure of propulsion efficiency, is defined as Isp=Fm˙g0I_{sp} = \frac{F}{\dot{m} g_0}, where FF is thrust, m˙\dot{m} is the total mass flow rate, and g0g_0 is standard gravity (9.80665 m/s²). In tripropellant systems, elevated Isp derives from higher exhaust velocity vev_e, governed by ve=2γRTcγ−1(1−(pepc)γ−1γ)v_e = \sqrt{\frac{2 \gamma R T_c}{\gamma - 1} \left(1 - \left(\frac{p_e}{p_c}\right)^{\frac{\gamma - 1}{\gamma}}\right)}
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