Liquid air cycle engine

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A Liquid Air Cycle Engine (LACE) is a type of spacecraft propulsion engine that attempts to increase its efficiency by gathering part of its oxidizer from the atmosphere. In a LOX/LH2 bipropellant rocket the liquid oxygen needed for combustion is the majority of the weight of the spacecraft on lift-off, so if some of this can be collected from the air on the way, it might dramatically lower the take-off weight of the spacecraft.

LACE was studied to some extent in the United States of America during the late 1950s and early 1960s, and by late 1960 Marquardt had a testbed system running. However, as NASA moved to ballistic capsules during Project Mercury, funding for research into winged vehicles slowly disappeared, and LACE work along with it.


Principle of operation

Conceptually LACE works by compressing and then quickly liquefying the air. Compression is achieved through the ram-air effect in an intake similar to that found on a high-speed aircraft like Concorde, where intake ramps create shock waves that compress the air. The LACE design then blows the compressed air over a heat exchanger, in which the liquid hydrogen fuel is flowing. This rapidly cools the air, and the various constituents quickly liquefy. By careful mechanical arrangement the liquid oxygen can be removed from the other parts of the air, notably water, nitrogen and carbon dioxide, at which point it can be fed into the engine as usual. The hydrogen is so much lighter than oxygen that the now-warm hydrogen is often dumped overboard instead of being re-used as fuel, at a net gain.

The use of a winged launch vehicle allows using lift rather than thrust to overcome gravity, which greatly reduces gravity losses.

One issue with the LACE system is that in order to appreciably reduce the mass of the oxygen carried at launch, a LACE vehicle needs to spend more time in the lower atmosphere to collect enough oxygen to supply the engines. This leads to greatly increased vehicle heating and drag losses, which therefore increases fuel consumption to offset the drag losses and the additional mass of the thermal protection system. This increased fuel consumption offsets somewhat the savings in oxidizer mass; these losses are in turn offset by the higher Isp (Specific impulse) of the air-breathing engine. Thus, the engineering trade-offs involved are quite complex, and highly sensitive to the design assumptions made.[1]

Other issues are introduced by the relative material and logistical properties of LOX versus LH2. LOX is quite cheap; LH2 is nearly two orders of magnitude more expensive.[2] LOX is dense (1.141 kg/L), whereas LH2 has a very low density (0.0678 kg/L) and is therefore very bulky. (The extreme bulkiness of the LH2 tankage tends to increase vehicle drag by increasing the vehicle's frontal area.) Finally, LOX tanks are relatively lightweight and fairly cheap, while the deep cryogenic nature and extreme physical properties of LH2 mandate that LH2 tanks and plumbing must be large and use heavy, expensive, exotic materials and insulation. Hence, much as the costs of using LH2 rather than a hydrocarbon fuel may well outweigh the Isp benefit of using LH2 in a Single-Stage-To-Orbit rocket, the costs of using more LH2 as a propellant and air-liquefaction coolant in LACE may well outweigh the benefits gained by not needing to carry as much LOX on-board.

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