The RS 68 engine

From 1990 onwards the United States has been developing the Evolved Expendable Launch Vehicle, a complementary vehicle to the Shuttle. The Delta family of launchers is one manifestation of this programme, and amongst its technological innovations has been the RS 68 engine (Figure 3.10), claimed to be the first new large rocket engine to be developed in the United States since the SSME. Its main features, compared with the SSME, are its simplicity and low cost. The number of separate components has been reduced by 80%, compared with the SSME, and the amount of manual manufacture has been cut to the minimum, most components being made by digitally controlled machines. The RS 68 has now been flight qualified on launches of the Delta IV vehicle. It is the United States counterpart to the Vulcain 2 engine on Ariane 5. The vacuum thrust is about twice that of the SSME, being 3.13 MN, while the exhaust velocity is relatively low for a liquid hydrogen-liquid oxygen engine, at 4,100 m/s. This is because of the low expansion ratio; this engine is intended to

Figure 3.10. The RS 68 engine firing. This is the expendable equivalent to the SSME, it is much cheaper to build, and has twice the thrust, all useful cost saving properties—only one engine needed rather than two. Courtesy NASA.

operate on the main stage of the Delta IV, and so is not optimised for vacuum. The sea-level thrust is relatively high at 2.89 MN, reflecting its purpose as an all-altitude booster. The weight of the engine is 6.6 tonnes, heavier than the SSME, but the thrust-to-weight ratio is about the same. Like the SSME, it can be throttled from 100% down to 60%. An engine of this thrust needs to make use of the gas generatorturbo-pump propellant delivery system to provide the necessary mass flow rates, and this contributes to the lower exhaust velocity; the hydrogen emerging from the turbo-pump exhaust is used for the roll-control thrusters of the Delta vehicle. Fundamentally, this is a low-cost expendable engine designed to provide high thrust for a heavy launcher.

3.5.8 The RL 10 engine

This engine, still a workhorse of the United States programme, has a heritage going back to the earliest liquid hydrogen-liquid oxygen engines designed in the United States (Figure 3.11); the first RL 10 was built in 1959. A pair of RL 10s power the Centaur upper stage, used on Atlas and Titan launchers. In its latest manifestation, the RL 10A-4-1, it has a vacuum thrust of 99 kN, weighs only 168 kg, and develops an exhaust velocity of 4,510 m/s. It is the archetypal upper-stage engine, optimised for vacuum use. It uses the expander cycle, with hydrogen heated in the cooling channels of the combustion chamber and upper nozzle powering the turbine of the liquid hydrogen pump, before entering the combustion chamber as gas. The liquid-oxygen pump is driven by a gear chain, from the hydrogen turbine; it delivers oxygen, as a liquid, to the injector. The engine is re-startable, giving a greater range of potential orbits.

Figure 3.11. An early photograph the RL 10 engine. The nozzle extension has been removed here. This engine is used in pairs to power the Centaur cryogenic upper stage, and has a heritage going back to the earliest use of liquid hydrogen and liquid oxygen in the United States.

Courtesy NASA.

Figure 3.11. An early photograph the RL 10 engine. The nozzle extension has been removed here. This engine is used in pairs to power the Centaur cryogenic upper stage, and has a heritage going back to the earliest use of liquid hydrogen and liquid oxygen in the United States.

Courtesy NASA.

The RL 10 engine has recently been considered as a potential chemical engine for Mars exploration, because it can be adapted to run using methane, instead of hydrogen, with the liquid oxygen. It is thought possible to produce methane on Mars from the carbon dioxide in the atmosphere, and this could be used for a return journey. Methane has other useful properties in that it is easy to store and has a high density as a liquid. It may therefore be the propellant of choice for long chemically propelled voyages. The exhaust velocity is of course smaller because of the presence of carbon dioxide in the combustion products; values as high as 3,700 m/s are predicted.

3.6 COMBUSTION AND THE CHOICE OF PROPELLANTS

Having examined the practicalities of propellant distribution in the liquid-fuelled engine, we shall now discuss the different types of propellant and the combustion process. Referring for the moment to Chapter 2, we recall that the exhaust velocity and thrust are related to the two coefficients c*, the characteristic velocity, and CF, the thrust coefficient. The thrust coefficient is dependent on the properties of the nozzle, while the characteristic velocity depends on the properties of the propellant and the combustion. It is defined by

For a given rocket engine the performance depends on the value of c *, defined above in terms of the molecular weight, the combustion temperature, and the ratio of specific heats, all referring to the exhaust gas. Different propellant combinations will produce different combustion temperatures and molecular weights. The exhaust velocity will also depend on the nozzle and ambient properties, but the primary factor is the propellant combination.

3.6.1 Combustion temperature

The exhaust velocity and thrust depend on the square root of the combustion temperature. The temperature itself varies a little depending on the expansion conditions, but the main dependence is on the chemical energy released by the reaction: the more energetic the reaction, the higher the temperature. Table 3.1 shows the combustion temperature under standard conditions for a number of propellant combinations.

The data in Table 3.1, which are calculated for adiabatic conditions, provide an insight into the effects of chemical energy. The combustion temperatures directly

The exhaust velocity and thrust defined by

Table 3.1. Combustion temperature and exhaust velocity for different propellants.

Oxidant Fuel

c Ve

O2 O2

2,386 4,550

1,783 3,580

2,530 4,790

1,724 3,420

1,731 3,420

N2O4 N2O4

(1) RP1 is a hydrocarbon fuel with hydrogen/carbon ratio 1.96, and density 0.81.

(2) MMH is monomethyl hydrazine.

(3) UDMH is unsymmetrical dimethyl hydrazine.

(4) The mixture ratios are optimised for expansion from 6.8 bar to vacuum.

reflect the chemical energy in the reaction. With oxygen as the oxidant, hydrogen produces a lower temperature than the hydrocarbon fuel RP1, the molecules of which contain more chemical energy. Fluorine and hydrogen produce a still higher temperature. This combination produces the highest temperature of any bi-propellant system. The corrosive nature of fluorine has prevented its use except with experimental rockets.

If the temperature is calculated theoretically for the complete reaction—for example, the combustion 2H2 + O2 = 2H2O—then a much higher value of about 5,000 K is predicted. In fact at this temperature, and for pressures prevalent in combustion chambers, much of the water formed by the reaction dissociates and absorbs some energy, lowering the temperature to the values shown in Table 3.1. If we deliberately introduce additional fuel, which cannot be burned without additional oxygen, then these atoms have to be heated by the same amount of chemical energy, and the temperature will be lowered further. This was discussed in the section on gas generators. Dissociation is an important phenomenon because it alters the molecular weight of the exhaust gases and the value of 7. For the oxygen-hydrogen combination the composition of the exhaust at 3,429 K is roughly 57% water and 36% hydrogen, with 3% monatomic hydrogen, 2% OH and 1% monatomic oxygen. The ratio of specific heats (7) for this mixture is about 1.25.

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