3.0 Hardware (Booster) Selection


3.1 Final Mass Estimate and Scope of Design

After several comparison calculations, the conclusion was reached that the Space Shuttle's greater lift capacity to LEO would be vital for a mission that would most likely require every kilogram of mass available.   However, for reasons discussed in 3.4 Launch System, this decision was not reached without carefully weighing a spectrum of safety and political concerns.  

Final Design - TK Model ms_sts.tk

Mission Time Event Mass
T+0:00 Shuttle Launch
2,040,000 kg
T+45:00 min LEO Delivery of Upper Stage
23,640 kg
T+2:00:00 hr Interplanetary Injection DV = 6.31 km/s
2,050 kg
T+2.73 year Powered Flyby (Semi-circularizing) Burn DV = 0.82 km/s
1,530 kg
T+2.73 year Heliocentric Z Component Addition Burn DV = 1.14 km/s
860 kg

The reader should keep in mind that this is a feasibility study.   Such issues as component mating, bus structure sizing, day level window specifics, and actual mission procedures have NOT been addressed.   They would be reserved for the more detailed analysis of a full scale design review.


3.2 Jovian Powered Flyby Booster

For the Jovian encounter phase of the mission, a Thiokol STAR 48V booster was selected for its low structure mass ratio (nominally 151 kg / 2,161 kg = 7.0 %).   The Isp of 292 seconds is on par with the field of booster examined (including the PAM-DII and IUS).   A significant factor in the selection of the Star 48V was the proven record of the 48 series.  

Also, given the exponential nature of the rocket equation, to minimize the mass needed at LEO delivery the most desirable spacecraft configuration would be all payload and fuel.   Since this is not possible, the 151 kg structure of the S48V was accepted as a penalty for a high momentum exchange with relatively low mass needed at Jupiter.   The S48V will be required to deliver the 0.82 km/s DV needed for eccentricity reduction at Jovian perigee and for the 1.14 km/s DV Z component velocity addition which, coupled with a z component slingshot about Jupiter, will give each transponder an inclination of 5.109º.


3.3 Upper Stage Booster

Considerations for the LEO-Interplanetary booster were slightly different than those for the Jovian encounter booster.   In this case, size and mass fraction were not as important as were raw throwing capability to get the 2,050 kg payload inserted into an interplanetary Hohmann transfer to Jupiter.   Two motors were considered: Martin Marietta's Centaur and TOS boosters.   The TOS lacked both the performance and mass to deliver a reasonably sized system to Jupiter.   The Centaur was more efficient and larger and was selected despite considerations discussed in the following chapter.

Booster Comparison
Isp mtotal mstruct Payload to Jupiter
TOS 294 seconds 10,800 kg 1,090 kg 137 kg
Centaur 444 seconds 18,910 kg* 3,540 kg 2,050 kg
*less than maximum capacity

3.4 Launch System

As mentioned in Section 2 the Space Shuttle and Titan IV launch systems were under consideration for the deployment of P-STAR.   Yet the STS LEO delivery maximum is 2,760 kg greater than that of the Titan IV (for 28º 110 NM orbits).   Applying a S48V flyby booster, the Centaur Isp, and the required DV for a Jovian Hohmann to the rocket equation this difference yields a 320 kg reduction (down to 540 kg) in transponder mass.   At first glance the choice of launch system may seem to obviously be the Shuttle.   Yet there has been a reluctance to carry Centaur boosted payloads in the STS fleet since the 1986 Shuttle disaster.   Faced with this problem two factors eventually overrode any hesitation to select a Centaur/Shuttle mix.  

First, there was a desire to stay within the realm of U.S. launch systems.   This was not for any jingoistic reasoning, since the Russian Energia system has certain tantalizing characteristics, but to keep from clouding the feasibility of the P-STAR system with the fesibility of multi-national launch cooperation which has yet to undergoe the test of Space Station construction.   Second, a system such as P-STAR would require a major restructuring of how the business of space navigation is carried out anyway.   In the face of such operational changes, the STS launch policy pales by comparison.


The Rocket Equation (solved for accelerated masses):
mstructure + mpayload = mfuel / ( eDV/(Isp*g) - 1 )


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Tim Crain
Graduate Aerospace Engineering
The University of Texas at Austin
crain@csr.utexas.edu
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