ESA Document - Emits - ESA
ESA Document - Emits - ESA ESA Document - Emits - ESA
s HMM Assessment Study Report: CDF-20(A) February 2004 page 80 of 422 Table 2-28 summarizes the dependencies of the slow abort. Note that the abort cost, which is within the mission manoeuvre budget, remains independent of the abort decision date, because the abort manoeuvre is best performed close to the arrival at Mars. A significant improvement of the efficiency could be achieved by allowing a powered swing-by. This is however not regarded here for the sake of obtaining conservative results. At any rate, abort could probably be performed even shortly prior to the arrival at Mars. The total transfer duration, with 904 days, remains within the nominal mission duration. Nevertheless, the time spent in deep-space increases, and so does the radiation dose due to the GCR, although the limits are not exceeded. The problem again is the hyperbolic arrival velocity, which is about 2 km/s higher than the nominal value. This translates into an increase of the entry velocity of the order of 0.8 km/s for the mission opportunity of 2033. The heat shield of the crew entry capsule was designed for an entry velocity of 12.5 km/s, which is not exceeded in this case. 2.7.12.1.4 Recovery from failed MOI If MOI fails completely due to a malfunction in the propulsion system, the spacecraft will perform a swing-by at Mars and end up in a heliocentric orbit much different from that of the Earth-Mars transfer. This orbit does not intersect that of the Earth. A natural close encounter with the Earth occurs in 2038. A correction manoeuvre of 2.5 km/s would enable a arrival in June 2038, five years after departure from the Earth. To fulfil the requirements of this abort scenario, major modifications would be required in the design of the THM. This analysis has not been carried out. Figure 2-40: Conditions in case of failed MOI Forcing an Earth encounter in June 2035 would require a correction manoeuvre of 5 km/s (available ∆V capability is 5.2 km/s). Arrival would occur at a hyperbolic velocity of 4.9 km/s, considerably faster than the nominal value. In this case the arrival velocity is not a problem, as the ERC is designed to cope with velocities up to 5.8 km/s. 2.7.12.1.5 Abort after MOI There is a theoretical option for a return to Earth via a Venus swing-by in December 2033 (i.e., a few weeks after nominal arrival). The scenario involves the following characteristics and cost:
s HMM Assessment Study Report: CDF-20(A) February 2004 page 81 of 422 • Mars escape: 3 December 2033 • Hyperbolic escape velocity: 5.94 km/s • TVI from 500 km orbit: 4.25 km/s, from HEO at least 3 km/s • Mars-Venus-transfer: 179 days • Venus swing-by minimum altitude: 6600 km • Venus-Earth transfer: 188 days • Earth arrival date: 5 December 2034 (around one year earlier than nominal) • Total transfer duration: 367 days • Earth arrival hyperbolic velocity: 4.74 km/s • No deterministic midcourse manoeuvres The total transfer duration is only one year, so arrival would occur one year earlier than for the nominal mission. The hyperbolic Earth arrival velocity is larger than in the nominal case but smaller than for the slow abort during Mars transfer (see above). The problem is the high escape velocity required from Mars. Insertion into the Venus transfer from the 500 km final orbit around Mars would cost 4.25 km/s, from the initial HEO at least 3 km/s. This value applies only if the HEO is oriented exactly as required for the escape, which is not the case. The escape manoeuvre from HEO would therefore incur a further large penalty. In the case of LMO, to provide the system with the required ∆V for the TVI, more propulsion modules should be added to the TEI, increasing the total mass to LEO by 1000 tonnes. In the case of HEO, the required ∆V can be provided by the second stage of the MOI plus the TEI without any mass penalty. The only other option appears to be to cancel the Mars landing and wait in orbit around Mars for the nominal return window in May 2035. 2.7.12.1.6 Abort after TEI No option was found for abort after TEI. 2.7.12.2 Options In addition to the regarded cases, there is a huge variety of other scenarios for which abort and recovery strategies need to be analysed. These relate primarily to the cases where critical manoeuvres are not fully executed due to a failure. Some examples are: • Incomplete execution of the TMI burn. A distinction must be made between two cases: - Failure while still in Earth orbit: This implies that the burn completed so far was not yet sufficient to inject into hyperbolic escape. A decision on whether to proceed with the burn at the next perigee pass, whether to abort and return the crew to Earth or to proceed in any other way must be made based on the gravity of the failure, the chances for recovery and the orbit achieved so far. - Failure after achieving escape: This implies that a hyperbolic orbit was achieved. Depending on the time of failure, the resulting heliocentric orbit will range from very close to the Earth orbit to very close to the nominal Mars transfer. Again, the abort/recovery options depend on the spacecraft and orbital conditions achieved. • Incomplete execution of MOI: Again a case distinction is necessary: - Failure after reaching a bound orbit: The burn removed enough orbital energy to insert into a possibly quite wide elliptical orbit around Mars. The reaction must again depend on the case and could consist of immediate stabilization of the orbit
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s<br />
HMM<br />
Assessment Study<br />
Report: CDF-20(A)<br />
February 2004<br />
page 80 of 422<br />
Table 2-28 summarizes the dependencies of the slow abort. Note that the abort cost, which is<br />
within the mission manoeuvre budget, remains independent of the abort decision date, because<br />
the abort manoeuvre is best performed close to the arrival at Mars. A significant improvement of<br />
the efficiency could be achieved by allowing a powered swing-by. This is however not regarded<br />
here for the sake of obtaining conservative results. At any rate, abort could probably be<br />
performed even shortly prior to the arrival at Mars.<br />
The total transfer duration, with 904 days, remains within the nominal mission duration.<br />
Nevertheless, the time spent in deep-space increases, and so does the radiation dose due to the<br />
GCR, although the limits are not exceeded.<br />
The problem again is the hyperbolic arrival velocity, which is about 2 km/s higher than the<br />
nominal value. This translates into an increase of the entry velocity of the order of 0.8 km/s for<br />
the mission opportunity of 2033. The heat shield of the crew entry capsule was designed for an<br />
entry velocity of 12.5 km/s, which is not exceeded in this case.<br />
2.7.12.1.4 Recovery from failed MOI<br />
If MOI fails completely due to a malfunction in the propulsion system, the spacecraft will<br />
perform a swing-by at Mars and end up in a heliocentric orbit much different from that of the<br />
Earth-Mars transfer. This orbit does not intersect that of the Earth. A natural close encounter<br />
with the Earth occurs in 2038. A correction manoeuvre of 2.5 km/s would enable a arrival in<br />
June 2038, five years after departure from the Earth.<br />
To fulfil the requirements of this abort scenario, major modifications would be required in the<br />
design of the THM. This analysis has not been carried out.<br />
Figure 2-40: Conditions in case of failed MOI<br />
Forcing an Earth encounter in June 2035 would require a correction manoeuvre of 5 km/s<br />
(available ∆V capability is 5.2 km/s). Arrival would occur at a hyperbolic velocity of 4.9 km/s,<br />
considerably faster than the nominal value. In this case the arrival velocity is not a problem, as<br />
the ERC is designed to cope with velocities up to 5.8 km/s.<br />
2.7.12.1.5 Abort after MOI<br />
There is a theoretical option for a return to Earth via a Venus swing-by in December 2033 (i.e., a<br />
few weeks after nominal arrival). The scenario involves the following characteristics and cost: