a UDMH = Unsymmetrical Dimethyl Hydrazine a UDMH = Unsymmetrical Dimethyl Hydrazine the weight of the oxidizer carried. As shown by Equation (4.12a), the launcher weight ratio is only a function of the carried oxidizer-to-fuel ratio and the weight ratio is determined by the propellant combination. From the propellant combinations in Table 4.5, the value of Wfuel/OWE for the different rocket propellant combinations was calculated and given in Table 4.6. Note that hydrogen carries the least fuel per OWE. With an oxidizer-to-fuel ratio of 6, that means the propellant load is 7.3 times the OWE. The hydrocarbons are five times greater and with an oxidizer to fuel ratio about 2.35, the propellant load is 17 times the OWE. The hypergolic propellants propellant load ranges from 19 to 20 times OWE. From Table 4.6 you can see why one of the famous Russian rocket designers, V. P. Glushko, chose the room temperature liquid UDMH and N2O4 for Proton and the submarine-launched ballistic missiles.

The importance of this relationship is that with minimal information a reasonable estimate of the fuel and propellant weight compared with OWE is available. Hydrogen provides the least weight ratio to orbit. Because the density of hydrogen is low, the volume required is the greatest.

The weight ratio is decreasing because the oxidizer weight is decreasing as a direct result of the oxidizer-to-fuel ratio. So from Figure 4.23, using hydrogen fuel, an all-rocket engine can reach orbital speed and altitude with a weight ratio of 8.1. An airbreathing rocket (AB rocket) or KLIN cycle can do the same with a weight ratio about 5.5. A combined cycle rocket/scramjet with a weight ratio of 4.5 to 4.0, and an ACES has weight ratio of 3.0 or less. So an airbreathing launcher has the potential to reduce the mass ratio to orbit by 60%. It is clear that results in a significantly smaller launcher, both in weight and size, and presumably also less expensive. To achieve this operationally, the design goal must be, "reduce the carried oxidizer''. It is more difficult if not impossible to achieve this progression of propulsion systems with fuels other than hydrogen. Methane is a cryogenic fuel, but it does not have the thermal capacity to liquefy or deeply cool air, so the hydrocarbon equivalent of a LACE or deeply cooled cycle is not possible. Ramjet/scramjet engine are possible with most of the liquid fuels, although for hydrocarbons the decomposition into free carbon will limit the temperature, and therefore the maximum speed is limited by the hydrocarbon thermal decomposition.

Examining the operational regions for each cycle concept we can make several observations.

(1) Chemical rocket, air augmented rocket and ram rocket maintain essentially a constant oxidizer-to-fuel ratio, with the weight ratio to achieve orbit decreasing because of the increased thrust produced by the air ejector system. For a vehicle for a rocket OWE equal to 76 metric tons and assuming the OWE of other propulsion systems at 76 t (plus any differential weight for the propulsion system), the TOGW for the three systems is:

WR O/F TOGW Savingsa


Rocket 8.10

Air augmented rocket 7.50

Ram rocket 6.50

a With respect to an all-rocket SSTO launcher.

6.00 616t 0 6.00 616t 0 5.80 5431 731

For the same liftoff weight of 616 t the payload for the three systems is 7.0, 6.0, and 15.4 tons respectively. As is usually the case for the air augmented rocket, the increased system weight is not offset by the increase in thrust unless the oxygen in the secondary air is burned. For the ram rocket the payload is more than doubled. The ram rocket is not any kind of technology challenge, as many afterburning turbojet engines have ejector nozzles (such as the mentioned Saab J-35 Viggen). The ram rocket is a simple way to increase payload to orbit using the same rocket engine, or to reduce the size and cost of the rocket engines for a fixed payload.

(2) LACE rocket, deeply cooled rocket and cooled turbojet-rocket (KLIN cycle) are other propulsion system concepts that build onto the basic rocket engine for increased performance. This propulsion system creates an airbreathing rocket operating to about Mach 5.5. All of the hardware required for the thermo-dynamic processing of the air has been built in one form or another over the last 45 years. No differentiation in weight is made for the liquid air cycle versus the deeply cooled. Historical data suggests that these two systems are essentially equal in total system weight. One of authors (PC) saw a 1 m3 liquid hydrogen/air heat exchanger operate for 1 min at Mitsubishi Heavy Industries in 1988 at outside air conditions of 38°C and 90% relative humidity without any water condensation on the heat exchanger tubes. The runtime was short because the container capturing the liquid air was overflowing and running down the ramp. So again this is not a technology issue, but (rather disappointingly) simply a decision-to-proceed issue. The KLIN cycle has the advantage of thrust for landing without the operation of a heat exchanger to provide the rocket with airbreathing capability. For a rocket vehicle with OWE equal to 76 metric tons and assuming the same OWE for other propulsion systems plus any system-specific differential, the TOGW for the two systems is:


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