1. The biconic vehicle is sized for injection into a trans-Mars trajectory by a Delta II class booster. It has an initial m = 950 kg and is shown on Fig.l. The vehicle characteristics are listed in Table 1. It has a hypersonic L/D ratio of 0.78 at (X= 25 deg in the 40-60 km range of altitudes, where most of aerodynamic deceleration accures. Variation of it's L/D with the angle of the attack is represented in Fig. 2. The angle of attack providing a given L/D ratio is maintained due to an offset of the vehicle's CG from the longitudinal axis. Control of the vehicle is realized through bank angle modulation. For simulation in POST a finite-roll-rate guidance scheme was used. This scheme places no constraints on either when the bank angle maneuvers are performed, the magnitude of bank angle changes, or the bank angle rates (except that the rates cannot exceed +11 deg/sec). Such a limit on the rate of change of bank angle is believed to be realistic for the chosen aerobraking vehicle with its size and mass characteristics. The vehicle also has a main liquid rocket engine, with 7sp= 320 sec, used for post-aerocapture propulsive burns to insert the vehicle into the target orbit. To compare aerocapture characteristics of the proposed biconic vehicle with a lower L/D configuration, a Viking type of aerobraking vehicle was chosen. This vehicle has similar mass-dimensional characteriristics to those of biconic, however, due to different shape it has an L/D ratio of 0.18 at a trim angle of attack of 11.2 deg. It was assumed that the Viking type of aerobraking vehicle utilises the same finite-roll-rate guidance scheme and that it has the same propulsion unit as the biconic vehicle.
2. One of the well-suited missions for the aerobraking vehicle described above can be the delivery to near Mars orbit and deployment on it's surface of a global network of miniature landers, which could provide continious information about Martian environment. For this purpose, the vehicle must be placed in highly elliptical, near-polar orbit with a period equal to that of one-half Martian day. This kind of orbit allows the deployment of a set of landers across Martian globe with relatively low deorbiting propulsive burns. This orbit also will enable the orbiting vehicle to communicate twice a day with the deployed landers in order to receive data on a regular basis. For a periapse radius of 3875 km and an apoapse radius of 21862 km such an orbit has an e = 0.6988. The flight profile begins at the entry interface, considered to be at 120 km altitude. The second phase is the aerobraking pass itself , during which the vehicle is targeted to the desired exit conditions, through bank angle modulation. The required exit conditions were (1): an minimum apoapse altitude of 200 km and (2): i = 93.125 deg. The next phase consisted of coasting along the resulting elliptical orbit with a subsequent rocket burn, to raise the periapse of the orbit to 500 km above the Martian surface, in order to stabilize it. In many cases, a combination of entry angle and bank angle modulations could be found that would produce a post-aerocapture trajectory with an apoapse radius equal to that of the one-half-sol orbit. For some cases, however, no such combination of entry angle and bank angles could be found and a last phase with a second rocket burn, at periapse, was performed, in order to adjust the resulting post-exit apoapse to the altitude of the onehalf-sol orbit.
3. Each mission opportunity resulted in different entry velocities, geographical locations of entry ( Table 2) and nominal atmospheric density profiles ( Fig.3). The 2001 mission has a significantly higher entry velocity
4. O.OE+00 2001 2003 2005 2005 0.78 0.78 0.78 0.18 6.226 5.636 5.520 5.520
5. 92 227.51 18.91 18.39