Each reactor concept was evaluated during mid-1990 by the Mission Analysis Panel to assess mission benefit figures of merit.(1) Results of this analysis were factored into the deliberations of the Technology Capabilities Panel. The initial trade space analysis of this Panel focused on the relationship between IMLEO and mission trip time for various reactor classes, using the example of the 2016 flight opportunity. Typical solid core configurations yielded trip times of 350 to 400 days at IMLEOs of 500 to 1,000 tons, with IMLEO quickly rising to over 3,000 tons for trip times of 300 days. For comparable IMLEO, the Low Pressure reactor offered reductions in trip times of 100 days for 600 tons IMLEO to 50 days at 2,000 tons IMLEO. Gas core reactors provided trip times as low as 200 days for 500 ton IMLEO, and as low as 150 day trip time at 600 tons.
Analysis of other launch opportunities produced triptimes that differed on average by as much as 150 days for comparable IMLEO, depending on mission mode and launch windows. A solid core reactor used during the significantly less favorable 2024 window requires an IMLEO of between 800 and 1,600 tons for a 450 day trip time, and 1,400 to 2,800 tons for a 400 day trip. And a 500 ton IMLEO gas core system would permit a triptime of from 250 to 400 days, depending on reactor performance.
The Mission Analysis Panel also considered the implications of multiple perigee burns. The baseline mission profile envisioned a three-periapsis (low point in orbit) burn Mars injection profile, in which the nuclear rocket would be fired on three consecutive perigee passes in Earth orbit, with each burn raising the vehicle's apogee, and the final burn placing the vehicle on a trans-Mars trajectory. This profile offered significant performance improvements over two-burn and single-burn profiles, with the greatest advantages for vehicles with low thrust-to-weight ratio engines. Penalties for single burn injections can range from about 5% for systems with engine thrust-to-weight ratio as high as 10, to about
10% for engines with thrust-to-weight ratios of 5. Drawbacks of the multiple-burn trajectory include hazards created by multiple passes through the Van Allen radiation belts, the danger of hypervelocity debris impact, and the increased standards of operational reliability and safety imposed by close-Earth flyby of the operating reactor.
The Mission Analysis Panel also considered mission trajectory profiles optimized for all-propulsive mission modes. These trajectories accepted higher trans-Mars injection velocities than were used in chemical+aerobrake mission analyses (14 km/sec versus 10.3 km/sec), to achieve lower Mars arrival velocities (5.3 km/sec versus 6.8 km/sec). This improved trajectory shaping resulted in a reduction in baseline IMLEO from 735 tons to 613 tons, although this resulted in the loss of free-return flight options.
NASA Lewis conducted a number of mission analysis studies to evaluate the potential contribution of NTR to Mars exploration.(2) One analysis found major benefits from using advanced nuclear propulsion with separate Piloted and Cargo vehicles supporting a 20 day stay on Mars, based on departure in 2002, using the following parameters:
Expedition 2002 Evolutionary 2004
System Propellant Isp IMLEO % IMLEO % tons Chem tons Chem
Chemical All Propulsive LOX/LH2 480 1782 260% 3800 663%
Chemical + Aerobrake LOX/LH2 480 686 100% 573 100%
1972 NTR LH2 800 675 98% 1133 198%
1989 NTR LH2 900 597 87% 1031 180%
Advanced NTR LH2 1000 496 72% 787 137%
NTR + Aerobrake LH2 900 421 61% 380 66%
Sensitivity analyses and trade studies were performed examining alternate launch windows in 2004, 2007 and 2010, as well as variations in the flight mode, including a traditional split/sprint and an all-up mission profile. These excursions did not result in significantly different results. Changing the reference mission to include recovery of the propulsion system in low Earth orbit at the conclusion of one mission for reuse on subsequent missions (the Evolutionary 2004 scenario) demonstrates significant disadvantages for nuclear concepts, other than the NTR + Aerobrake. This analysis concluded that:(3)
"Based on the ground-rules and assumptions used in this study, the chem/AB system generally requires somewhat less IMLEO that the NTR for evolutionary-type missions with this trend reversing for the expeditionary mission scenario. Overall, the differences are such that both systems can be considered comparable. Combining the NTR with an aerobrake results in large mass savings, however, this option requires the development of both technologies.
"The technology development and validation work for aerobrakes required to derive a more sophisticated and informed set of ground rules remains to be done. Consequently it is not clear at this time which technology has the lower IMLEO. The NTR, having an established technology base and substantial mass advantage over the all-chemical system, provides a credible all-propulsive option for piloted missions to Mars. The Chem/AB has comparable IMLEO requirements, but lacks the technology maturity to make this option as credible as the NTR."
NASA Marshall Space Flight Center (MSFC) conducted a mission analysis in conjunction with Boeing, which considered a very wide range of propulsion technologies and mission architectures. This analysis concluded that:(4)
"For the minimum activity level, cryogenic all-propulsive conjunction with storable Mars ascent and Earth return provides a minimum-cost, risk balanced program. Aerobraking with a Venus swingby opposition profile offers a shorter-duration (typically 500 days vs. 900 days) alternative... The NTR architecture can provide round trips to Mars of less than one year. For fast trips, a split-sprint mission profile reduces total mass in Earth orbit (IMLEO) by about a factor of 2... At a median activity level, NTR enables flexibility to fly conjunction or opposition as well as economic recovery of the Mars transfer and NTR vehicle. If there are more than three missions, development costs of NTR appear amortized. The NTR architecture is capable of fast Mars round trips, a year or less, using a split mission profile. However, these modes are entirely expendable and use up a lot of hardware. The indicated life-cycle penalty for fast trips over a Mars exploration program is many billions of dollars... Desirability of the direct Mars surface rendezvous mode is unresolved. It appears too risky (lack of abort modes) for an initial mission."
Boeing's primary interest in NTR is in the area of mission analysis and integration.(5) A recent review of the applicability of NTR mission applications noted that the NERVA design specifications, down to the subsystem component line drawing level, are still retrievable. A baseline Mars mission envisions a crew of 4 using a 925 second Isp 333 kN NERVA- Derived Reactor (NDR) flying an all-propulsive profile with crew Earth-return either using an Apollo-type ECCV, or propulsive braking of the entire vehicle into a 500 Km perigee 24 hour orbit for subsequent vehicle reuse. This mission profile required a total trip-time of 434 days with a 30 day stay on Mars and a Venus gravity-assisted return, assuming the following parameters:
Mass
MTV 34,939 Kg (no artificial-G)
MEV 73,000 Kg (includes aerobrake)
ECCV 7,000 Kg (optional)
NERVA Engine 9,684 Kg
Shadow Shield 4,500 Kg
Propellant Tank Fraction 14 %
Cooldown Penalty 3 %
and the following timeline:
EVENT DATE Delta-V Mass
IMLEO 735,190 Kg
Earth Departure 25 Feb 2016 4,201 M/Sec - 329,238 Kg
Arrive Mars 31 Jul 2016 3,870 M/Sec - 177,252 Kg
Depart Mars 31 Aug 2016 3,900 M/Sec - 59,245 Kg
Venus encounter 10 Mar 2017
Earth Return 5 May 2017 2,653 M/Sec - 41,601 Kg
The vehicle for this mission underwent an iterative design process to optimize propellant tank configuration. These configurations incorporated separate propellant tanks for each main propulsive maneuver, which are jettisoned at the conclusion of the burn. The tanks are mounted on a truss structure, which was found to offer an order of magnitude mass advantage relative to using propellant tanks as structural members. Although the baseline vehicle did not provide artificial gravity for the crew, enlarging the truss structure from 7 to 14 standard SSF truss bays resulted in a 140 meter overall length vehicle that could be rotated to provide 1/3 G artificial gravity, with a mass penalty of only 16,000 Kg above the baseline. Addition of more trusses to lengthen the vehicle to 215 meters would provide full Earth gravity when rotated, with a total IMLEO of 787,300 kg, 52,000 kg over the baseline. This 7% mass penalty included heavier truss structures, added RCS and propellant for spin-up and spin-down for mid-course correction maneuvers, and de-spun joints for power and communications.
The baseline NERVA NTR was assumed to provide a thrust-to-weight ratio of 3.5, while a 3,401 Kg PBR providing a thrust to weight (T/W) ratio of 15 would reduce IMLEO from 735,190 Kg to 673,425 Kg. The analysis also considered use of high performance (1250 Isp) low thrust ( three 44 kN) NTR engines using three peripasis burns for Earth departure.
An earlier analysis, using a slightly lighter crew habitation module, considered the impact of departure year on IMLEO, and determined that the 2016 baseline had the highest IMLEO, but the shortest trip time.
YEAR Class IMLEO Duration Burntime
(Days) (Hours)
2010 Opposition 525,000 Kg 638 2.4
2013 Opposition 437,000 Kg 631 1.9
* 2016 Opposition 648,000 Kg 434 3.2
2018 Opposition 434,000 Kg 631 2.4
2020 Opposition 520,000 Kg 603 1.9
2023 Opposition 583,000 Kg 571 2.8
2023 Conjunction 368,000 Kg 970 1.4
2024 Opposition 509,000 Kg 618 2.3
Another Boeing design excursion considered a low energy 2005 conjunction mission with a 928 day trip time, including 482 days on Mars, and an ECCV for crew return, with the nuclear vehicle expended (rather than returned to Earth orbit). This 1/3 G artificial gravity mission configuration carried a crew of 6 and two MEVs of the type previously discussed (in contrast to the 4 crew members and single MEV of the baseline mission), and required a 605,300 kg IMLEO. Sensitivity analyses noted that each additional crew member above the initial 4 imposed a 17,000 Kg IMLEO penalty, while reuse of the Earth Arrival Stage imposed a 16,000 Kg penalty. Conducting the mission without artificial gravity was found to yield an IMLEO of 594,000 Kg, while full Earth artificial gravity could impose an IMLEO penalty of from 40,000 Kg to 56,000 Kg, depending on whether cryogenic or storable RCS propellants were used, respectively.
This analysis also considered several Lunar NTR vehicle configurations, including a single lander version which would perform the same mission as the chemical LTV of the NASA 90-day study; and a larger vehicle carrying two landers. Design excursions also considered use of a PBR with a T/W ratio of 15 instead of the 3.5 T/W NDR.
Landers NDR (3.5 T/W) PBR (15 T/W)
1 197,000 Kg 170,000 Kg
2 293,000 Kg 267,000 Kg
NTR Advantages include high performance without use of aerobrake, LTV and LEV reusability, and 90% propulsion system commonality with NTR Mars vehicles (with the exception of propellant tank size). The Lunar and Martian propulsion profiles were regarded as strikingly similar, with the exceptions of propellant boiloff on longer Mars missions, and larger Delta-V requirements for Mars opposition-class missions. Total round-trip burn-times are on the order of 45 minutes, resulting in small fission-product buildup, with the crew occupying the LTV solar flare radiation shelter during reactor operations.
1. Wickenheiser, Tim, (NASA Lewis Research Center), "Mission Analysis Panel Report," NEP/NTP Workshop Feedback Meeting, (Hilton Nassau Bay Hilton, Houston, TX, 15 November 1990).
2. Borowski, Stanley, et al, "Performance Comparisons of Nuclear Thermal Rocket and Chemical Propulsion Systems for Piloted Missions to Phobos/Mars," Seventh Symposium on Space Nuclear Power Systems, Albuquerque, NM, 7-10 January 1990, pages 38-45.
3. Burowski, S.K., et al, "Performance Comparisons of Nuclear Thermal Rocket and Chemical Propulsion Systems for Piloted Missions to Phobos/Mars," 40th Congress of the International Astronautical Federation, Malaga, Spain, October 1989, paper IAF-89-027.
4. Adams, Alan, et al, "Overview of Mars Transportation Options and Issues," AIAA Space Programs and Technologies Conference, Huntsville, AL, 25-28 September 1990, AIAA-90-3795.
5. Donohue, B., "NTR Mission Applications," NASA Nuclear Propulsion Workshop, NASA Lewis Research Center, Cleveland OH, July 10-12, 1990.
Hornung, R.J., "Quick Trips to Mars," NASA Nuclear Propulsion Workshop, NASA Lewis Research Center, Cleveland OH, July 10-12, 1990.