JuanCastillo29/space-mission-intelligence-data huggingface dataset:
Enter control inputs, manually with gloved hand, to configure Earth Descent Vehicle (EDV) computers, for Earth Surface Descent.
Enter control inputs, manually with gloved hand, to adjust EDV pitch and yaw during Earth Surface Descent.
Enter control inputs, manually with gloved hand, to maneuver Earth Descent Vehicle (EDV) during Earth Surface Descent.
Enter control inputs, manually with gloved hand, to override automatic firing of retro rockets during Earth Surface Descent if automated system fails.
Enter control inputs, manually with gloved hand, to adjust/override automatic systems in response to system displays, if necessary.
Adjust retro rocket controls, manually with gloved hand during buffeted descent, to maneuver EDV to correct landing location, if necessary.
Adjust controls, manually with gloved hand, to shut down retro rocket and other systems after touch down on Earth surface.
Verify automated engine shut down by visually monitoring displays to ensure safing of Earth Descent Vehicle (EDV) after touch down.
Enter control inputs, manually with gloved hand, to activate engine shut down if automated shut down fails, to ensure safing of Earth Descent Vehicle (EDV).

Verify automated landing system shut down by visually monitoring displays to ensure safing of Earth Descent Vehicle (EDV) after touch down.
Activate landing system shut down manually by control inputs if displays indicate failure of automated shut down, to ensure safing of EDV after touch down.
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Growth Technology for “Fast Transit” Human
Missions to Mars

Stanley K. Borowski Glenn Research Center, Cleveland, Ohio David R. McCurdy and Thomas W. Packard Vantage Partners, LLC, Brook Park, Ohio
Prepared for the Space 2013 Conference and Exposition sponsored by the American Institute of Aeronautics and Astronautics San Diego, California, September 10–12, 2013
National Aeronautics and Space Administration
Glenn Research Center Cleveland, Ohio 44135
The authors wish to thank Chris Moore and John Warren from NASA Headquarters, Mike Houts from NASA Marshall Space Flight Center, along with John Taylor and Gary Kelm from NASA Glenn Research Center for their continued encouragement and strong support of this work. The author Stanley K. Borowski also expresses his thanks to Bob Sauls (bob.sauls@xp4d.com) for artwork and images depicted in Figures 5, 6, and 7 of this paper.
Level of Review: This material has been technically reviewed by technical management.
Available from
NASA Center for Aerospace Information 7115 Standard Drive Hanover, MD 21076–1320
National Technical Information Service 5301 Shawnee Road Alexandria, VA 22312
Available electronically at http://www.sti.nasa.gov Nuclear Thermal Propulsion (NTP):

A Proven, Growth Technology for “Fast Transit” Human Missions to Mars
Stanley K. Borowski National Aeronautics and Space Administration Glenn Research Center Cleveland, Ohio 44135

David R. McCurdy and Thomas W. Packard Vantage Partners, LLC Brook Park, Ohio 44142
The “fast conjunction” long surface stay mission option was selected for NASA’s recent Mars Design Reference Architecture (DRA) 5.0 study because it provided adequate time at Mars (~540 days) for the crew to explore the planet’s geological diversity while also reducing the “1-way” transit times to and from Mars to ~6 months. Short transit times are desirable in order to reduce the debilitating physiological effects on the human body that can result from prolonged exposure to the zero-gravity (0-gE) and radiation environments of space. Recent measurements from the RAD detector attached to the Curiosity rover indicate that astronauts would receive a radiation dose of ~0.66 Sv (~66 rem)—the limiting value established by NASA—during their 1-year journey in deep space. Proven nuclear thermal rocket (NTR) technology, with its high thrust and high specific impulse (Isp ~900 s), can cut 1-way transit times by as much as 50 percent by increasing the propellant capacity of the Mars transfer vehicle (MTV). No large technology scale-ups in engine size are required for these short transit missions either since the smallest engine tested during the Rover program—the 25 klbf “Pewee” engine is sufficient when used in a clustered arrangement of three to four engines. The “Copernicus” crewed MTV developed for DRA 5.0 is a 0-gE design consisting of three basic components: (1) the NTP stage (NTPS); (2) the crewed payload element; and (3) an integrated “saddle truss” and LH2 propellant drop tank assembly that connects the two elements. With a propellant capacity of ~190 t, Copernicus can support 1-way transit times ranging from ~150 to 220 days over the 15-year synodic cycle. The paper examines the impact on vehicle design of decreasing transit times for the 2033 mission opportunity. With a fourth “upgraded” SLS/HLV launch, an “in-line” LH2 tank element can be added to Copernicus allowing 1-way transit times of 130 days. To achieve 100 to 120 day transit times, Copernicus’ saddle truss/drop tank assembly is replaced by a “star truss” assembly with paired modular drop tanks to further increase the vehicle’s propellant capacity. The HLV launch count increases (from ~5 to 7) and a fourth engine is needed to reduce total mission burn time and gravity losses. Using a “split mission” approach, the NTPS, in-line tank and the saddle truss/LH2 drop tank elements can be configured as a pre-deployed Earth Return Vehicle/propellant tanker supporting 90-day crewed mission transits. The split mission approach also eliminates the need for onorbit assembly. Mission scenario descriptions, key features and operational characteristics for five different vehicle configurations are presented.
IMLEO initial mass in low Earth orbit K

temperature (degrees Kelvin) klbf

thrust (1000’s of pounds force) LEO

low Earth orbit (= 407 km circular) LOX/LH2 Liquid oxygen/liquid hydrogen propellant RAD

Radiation Assessment Detector carried on Mars Science Laboratory’s Curiosity rover SLS/HLV Space Launch System/Heavy Lift Vehicle Sv

Sievert (1 Sv = 100 rem) t

metric ton (1 t = 1000 kg) V

velocity change increment (km/s)
Introduction and Background

In 2007 and 2008, NASA conducted a multi-center, agency-wide study that examined the mission, payload and transportation system options and requirements needed to land humans on Mars in the post-2030 timeframe. The Mars Design Reference Architecture (DRA) 5.0 study (Ref. 1) provided an update to NASA’s earlier DRM 3.0 and 4.0 studies (Refs. 2 to 5) and was conducted in two phases. In Phase I, key architectural approaches, mission drivers and technology options were evaluated to support a downselection process. Selected architectural approaches and systems concepts were then further refined in Phase II while also evaluating different surface exploration strategies. A three mission set of Mars landing missions over an ~10 year period was assumed with each mission carrying a crew of six. Opposition and fast-conjunction trajectories supporting short and long surface stay landing missions were examined (shown in Fig. 1) during Phase I. Mission characteristics and features associated with each trajectory type were evaluated and compared against the in-space transportation system requirements and the amount of scientific investigation that could be achieved using each mission type. The scientific community overwhelmingly preferred going to different sites and staying as long as possible so the crew could sample the planet’s rich geological diversity and maximize the scientific return on each mission. While opposition trajectories typically have short round trip mission times, the stay times at Mars are also short. This trajectory class also usually has one short and one long transit leg with the latter involving a Venus swing-by (VSB) maneuver that can bring the spacecraft relatively close to the Sun. By contrast, fast conjunction trajectories have long round trip and surface stay times, but transit times to and from Mars can be relatively short and no VSB is required on either mission leg. The long surface stay mission also has a lower total V requirement than the short stay option, and therefore requires less propellant and less mass delivered to LEO. Because of its attractive features—lower total V, a long surface stay, short transit times, and no VSB—the fast conjunction/long stay mission option was selected as the baseline for DRA 5.0.

[Figure: Figure 1.—Features and characteristics associated with opposition and fast conjunction Mars missions.]
. 9 and 10) consisting of three basic components: (1) the NTPS; (2) the crewed payload element; and (3) an integrated “saddle truss” and LH2 propellant drop tank assembly that connects the two elements. With a propellant capacity of ~190 t in its two tanks, Copernicus can achieve 1-way transit times to Mars and back ranging from ~150 to 220 days over a 15-year synodic cycle starting in 2033.

[Figure: Figure 2.—“Copernicus” crewed MTV in LEO prior to departure.]

This paper examines the possibilities for achieving short transits times by reconfiguring the Copernicus design to allow it to carry larger quantities of LH2 propellant needed for these more demanding missions. The paper covers the following topic areas. First, the operational principles and performance characteristics of the baseline 25 klbf Pewee-class engine used in this analysis are presented along with a summary of the Rover/NERVA programs’ technical accomplishments. Key NTP development activities currently funded under NASA’s Advanced Exploration Systems (AES) program are also mentioned. Mission and transportation system ground rules and assumptions used in the analysis are then presented along with trajectory data for a fast conjunction 2033 mission that covers a range of transit times. Copernicus’ key components and mission capabilities are reviewed next. To reduce 1-way transit times down to 130 days, then to 100 days, modifications are made to the Copernicus configuration incrementally—first with the addition of an in-line tank, then with the addition of paired sets of drop tanks—to increase the vehicle’s propellant loading. A “split crewed mission” option that does not require orbital assembly and uses Copernicus’ components is also described that can support a crewed mission with 90-day transit times. Mission scenario descriptions, key features and characteristics for five different vehicle configurations are presented along with the engines’ operational requirements. The paper ends with a summary of our findings and some concluding remarks.
NTR System Description and Performance Characteristics

The NTR uses a compact fission reactor core containing 93 percent “enriched” Uranium (U)-235 fuel to generate 100’s of megawatts of thermal power (MWt) required to heat the LH2 propellant to high
exhaust temperatures for rocket thrust. In an “expander cycle” Rover/NERVA-type engine (Fig. 3), high pressure LH2 flowing from either a single or twin turbopump assembly (TPA) is split into two paths with the first cooling the engine’s nozzle, pressure vessel, neutron reflector, and control drums, and the second path cooling the engine’s tie-tube assemblies. The flows are then merged and the heated H2 gas is used to drive the TPAs. The hydrogen turbine exhaust is then routed back into the reactor pressure vessel and through the internal radiation shield and core support structure before entering the coolant channels in the reactor core’s fuel elements. Here it absorbs energy produced from the fission of U-235 atoms, is superheated to high exhaust temperatures (Tex ~2550 to 2950 K depending on fuel type and uranium loading), then expanded out a high area ratio nozzle (~300:1) for thrust generation. Controlling the NTR during its various operational phases (startup, full thrust and shutdown) is accomplished by matching the TPA-supplied LH2 flow to the reactor power level. Multiple control drums, located in the reflector region surrounding the reactor core, regulate the neutron population and reactor power level over the NTR’s operational lifetime. The internal neutron and gamma radiation shield, located within the engine’s pressure vessel, contains its own interior coolant channels. It is placed between the reactor core and key engine components to prevent excessive radiation heating and material damage.

[Figure: Figure 3.—Schematic of “expander cycle” NTR engine with dual LH2 turbopumps.]
Assumptions

The fast conjunction 2033 opportunity was selected as the baseline mission for this analysis since missions performed in the 2033 to 2035 timeframe have V budgets near their minimum values over the 15-year synodic cycle. As a result, it is expected that the shortest achievable transit times would be found in this opportunity. Table 2 shows the ideal V budgets as a function of decreasing transit time for the outbound Earth to Mars mission leg. It is assumed that the crewed MTV departs from a 407 km circular LEO, then captures into and departs from a 250 by 33,793 km elliptical Mars orbit (EMO) with a 24-hr orbital period. Gravity losses are also added to the ideal Vs and depend on the particular mission C3, vehicle thrust-to-weight (F/W), and propulsion system specific impulse.

[Table: TABLE 2.—TRAJECTORY DATA FOR 2033 FAST CONJUNCTION MISSION—OUTBOUND EARTH-TO-MARS LEG
TABLE 3.—TRAJECTORY DATA FOR 2033 FAST CONJUNCTION MISSION—MARS-TO-EARTH RETURN LEG]

[Table: Table 3 shows the ideal V budgets for the same range of transit times for the inbound Mars to Earth mission leg with Mars departure occurring in late 2035. The last column in the table shows the corresponding stay time at Mars assuming equal outbound and inbound transit times. As transit times decrease, it becomes necessary to perform a slowdown burn on the return leg to ensure that arrival Vinf at Earth is less than 6.81 km/s at an altitude of ~125 km. For 180-day transits to and from Mars and a 575-day stay time at Mars, the total round trip time and V budget (minus g-loss) for the mission are 935 days and 6.045 km/s, respectively. Cutting the transit times in half to 90-days increases the Mars stay time to ~673 days, but decreases the round trip time to ~853 days. The mission V budget however, increases by ~250 percent to ~15.175 km/s! While challenging, transit times this short are possible using NTP. The mission and transportation system ground rules and assumptions used in this paper are summarized in Tables 4 and 5, respectively. Table 4 provides information about the mission profiles and operational scenarios, and the assumed parking orbits at Earth and Mars. In addition to the large V requirements for the primary propulsion maneuvers—trans-Mars injection (TMI), Mars orbit capture (MOC), and trans-Earth injection (TEI)—smaller V maneuvers are needed for rendezvous and docking (R&D) of vehicle components during the LEO assembly phase, for spacecraft attitude control during inspace transit, and for “split crewed mission” orbital operations at Mars.]

[Table: TABLE 4.—MISSION AND PAYLOAD GROUND RULES AND ASSUMPTIONS
TABLE 5.—NTR TRANSPORTATION SYSTEM GROUND RULES AND ASSUMPTIONS]
drop tank and connects the NTR stage to the forward payload section. It is ~27.7 m long and has a mass of ~9 t. The saddle truss is open underneath allowing the drained LH2 drop tank to be jettisoned after the TMI burn is completed. The ~22.7 m long LH2 drop tank has a mass of ~21 t and a propellant capacity of ~102.4 t but is also off-loaded and only carries ~91.9 t of LH2 propellant.

[Figure: Figure 7.—Key features and component lengths of the Copernicus Mars transfer vehicle.]

Copernicus’ third and final component is its payload section. In DRA 5.0 it was designed for launch as a single integrated unit and thus determines the overall size of the PL envelope. The integrated payload element is ~33.8 m long and includes the short saddle truss, “T-shaped” DM and transfer tunnel, consumables container, TransHab and crew, and the Orion MPCV. The DM provides “secondary access” to Copernicus for the crew delivery MPCV and the MAV (see Figs. 6 and 7). Following the crew’s return from Mars and MAV separation, the DM and attached consumables container are both jettisoned to reduce vehicle mass prior to TEI (see Fig. 5). The total crewed payload mass at TMI is ~67.3 t distributed as follows: (1) short saddle truss (~5.1 t); (2) DM and transfer tunnel (~2.8 t); (3) contingency consumables and jettisonable container (~10.7 t); (4) TransHab with its primary PVAs (~27.5 t); (5) transit consumables (~4.6 t); (6) crew (~0.6 t); (7) MPCV (~10 t); and (8) forward RCS and propellant (~6 t). Copernicus’ total RCS propellant loading is ~11.9 t with the “post-TMI” RCS propellant load split between the NTPS (~7 t) and the short saddle truss forward cylindrical adaptor ring (~4.9 t). Lastly, for this particular mission, the performance requirements on operating time and restart for Copernicus’ three 25.1 klbf NTR engines are quite reasonable. For the round trip mission, there are four primary burns (three restarts) that use ~168.7 t of LH2 propellant. With ~75.3 klbf of total thrust and a Isp
of ~906 s, the total engine burn time for the mission is ~74.6 min (~53.2 min for the “2-perigee burn” TMI maneuver, ~13.6 min for MOC, and ~7.8 min for TEI), well under the ~2 hr accumulated engine burn time and 27 restarts demonstrated by the NERVA eXperimental Engine—the NRX-XE (Ref. 8).
Copernicus “Growth” Options Allowing Faster Transit Time Missions

As discussed in Section 4.0, Copernicus was sized so it can perform all fast conjunction missions in the 2033 to 2048 timeframe. For the 2033 mission opportunity, with its propellant tanks near their maximum capacity of ~190 t, Copernicus can achieve 150-day transit times. Besides human Mars missions, the “3-element” Copernicus spacecraft (Fig. 8) has significant capability that can be exploited for a variety of other mission applications such as a reusable human mission to the large, high energy near Earth asteroid—Apophis in 2028. “Scaled down” versions of Copernicus can also be used for reusable lunar cargo delivery and crewed lunar landing missions. Mission details and vehicle characteristics for these applications are provided elsewhere (Ref. 10). Figure 9 shows the round trip mission V versus transit time for the 2033 opportunity using the trajectory data found in Tables 2 and 3. As is evident, the V requirements rise dramatically as the transit times are shortened. Decreasing the transit times by 60 days—from 180 to 120 days—increases the total mission V by ~46 percent. A further 60-day reduction in transit time—down to 90 days each way— increases the total mission V by ~250 percent. To achieve transit times less than 150 days, it is necessary to modify the Copernicus spacecraft so it can carry additional propellant. This will be done in incremental steps—first with the addition of an inline tank, then with the addition of paired sets of drop tanks—to show what can be achieved, but also to show what is required in terms of extra hardware, engine operating time and additional HLV launches.

[Figure: Figure 8.—Growth paths and alternative missions for Copernicus spacecraft using modular components.]

[Figure: Figure 9.—Total mission V versus transit time.]

[Figure: Figure 10.—Copernicus-2 option: Supports 130-day transit times for the 2033 mission.]