NASA/TM--1998-208834/REV1 AIAA-98-3883

Vehicle and Mission Design Options for the

Human Exploration of Mars/Phobos Using

"Bimodal" NTR and LANTR Propulsion

Stanley K. Borowski and Leonard A. Dudzinski

Glenn Research Center, Cleveland, Ohio

Melissa L. McGuire

Analex Corporation, Brook Park, Ohio

December 2002
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NASA/TM--1998-208834/REV1 AIAA-98-3883

Vehicle and Mission Design Options for the

Human Exploration of Mars/Phobos Using

"Bimodal" NTR and LANTR Propulsion

Stanley K. Borowski and Leonard A. Dudzinski

Glenn Research Center, Cleveland, Ohio

Melissa L. McGuire

Analex Corporation, Brook Park, Ohio

Prepared for the

34th Joint Propulsion Conference

cosponsored by the AIAA, ASME, SAE, and ASEE

Cleveland, Ohio, July 13-15, 1998

National Aeronautics and

Space Administration

Glenn Research Center

December 2002
Acknowledgments

The authors wish to express their thanks to GRC management (Pat Symons, Harry Cikanek, and Joe Nieberding)

and NASA Headquarters (Lewis Peach) for support and encouragement during the course of this work, and to a

number of individuals for key contributions to various topics addressed in this stud_

They include: Don Culver (Aerojet) on bimodal CIS engine design issues, Lee Mason (NASA Glenn) on Brayton

cycle PCU analysis and system characterization, Dave Plachta (NASA Glenn) on LH 2thermal protection and active

refrigeration systems, Mike Stancati (Science Applications International Corporation (SAIC)) on disposal AV

estimates, and Pat Rawlings (SAIC) for artwork depicted in Figure 2.

Document Change History

This printing, numbered as NASA/TM--1998-208834/REV1, December 2002, replaces the previous version,

NASA/TM--1998-208834, December 1998.

Nontechnical changes have been made to the text and the figures have been recreated to improve the quality.

NASA Center for Aerospace Information

7121 Standard Drive

Hanover, MD 21076

Available from

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Available electronically at http://gltrs.grc.nasa.gov
VEHICLE AND MISSION DESIGN OPTIONS FOR THE HUMAN

EXPLORATION OF MARS/PHOBOS USING "BIMODAL" NTR

AND LANTR PROPULSION

Stanley K. Borowski* and Leonard A. Dudzinski**

National Aeronautics and Space Administration

Glenn Research Center

Cleveland, Ohio 44135

Phone: 216-977-7091 and 216-977-7107

Melissa L. McGuire**

Analex Corporation

Brook Park, Ohio 44142

216-977-7128

ABSTRACT

The nuclear thermal rocket (NTR) is one of

the leading propulsion options for future human

missions to Mars because of its high specific

impulse (Isp~850-1000 s) capability and its attractive engine thrust-to-weight ratio (~3-10). To stay

within the available mass and payload volume

limits of a "Magnum" heavy lift vehicle, a high

performance propulsion system is required for

trans-Mars injection (TMI). An expendable TMI

stage, powered by three 15 thousand pounds

force (klbf) NTR engines is currently under

consideration by NASA for its Design Reference

Mission (DRM). However, because of the

miniscule burnup of enriched uranium-235 during

the Earth departure phase (~10 grams out of

33 kilograms in each NTR core), disposal of the TMI

stage and its engines after a single use is a costly

and inefficient use of this high performance stage.

By reconfiguring the engines for both propulsive

thrust and modest power generation (referred to

as "bimodal" operation), a robust, multiple burn,

"power-rich" stage with propulsive Mars capture

and reuse capability is possible. A family of

modular "bimodal" NTR (BNTR) vehicles are

described which utilize a common "core" stage

powered by three 15 klbf BNTRs that produce

50 kWe of total electrical power for crew life support,

an active refrigeration / reliquification system

for long term, "zero-boiloff" liquid hydrogen (LH2)

storage, and high data rate communications.

An innovative, spine-like "saddle truss" design

connects the core stage and payload element

*Ph.D./NuclearEngineering,SeniorMemberAIAA

**AerospaceEngineer,MemberAIAA

and is open underneath to allow supplemental

"in-line" propellant tanks and contingency crew

consumables to be easily jettisoned to improve

vehicle performance. A "modified" DRM using

BNTR transfer vehicles requires fewer transportation system elements, reduces IMLEO and

mission risk, and simplifies space operations. By

taking the next logical step--use of the BNTR for

propulsive capture of all payload elements into

Mars orbit--the power available in Mars orbit grows

to 150 kWe compared to 30 kWe for the DRM.

Propulsive capture also eliminates the complex,

higher risk aerobraking and capture maneuver

which is replaced by a simpler reentry using a

standardized, lower mass "aerodescent" shell.

The attractiveness of the "all BNTR" option is

further increased by the substitution of the

lightweight, inflatable "TransHab" module in place

of the heavier, hard-shell hab module. Use of

TransHab introduces the potential for propulsive

recovery and reuse of the BNTR / Earth return

vehicle (ERV). It also allows the crew to travel to and

from Mars on the same BNTR transfer vehicle

thereby cutting the duration of the ERV mission in

half--from ~4.7 to 2.5 years. Finally, for difficult

Mars options, such as Phobos rendezvous and

sample return missions, volume (not mass) constraints limit the performance of the "all LH2" BNTR

stage. The use of "LOX-augmented" NTR (LANTR)

engines, operating at a modest oxygen-tohydrogen mixture ratio (MR) of 0.5, helps to

increase "bulk" propellant density and total thrust

during the TMI burn. On all subsequent burns, the

bimodal LANTR engines operate on LH2 only (MR=0)

to maximize vehicle performance while staying within

the lift capability of two Magnum launches.

NASA/TM-- 1998-208834/REV 1 1
INTRODUCTION AND BACKGROUND

The possible discovery of ancient microfossils

in the Mars meteorite ALH84001, along with the

excitement provided by the Mars Pathfinder and

current Mars Surveyor missions 1 has stirred

worldwide interest in the question of extraterrestrial life and in NASA's plans for future

human exploration missions to Mars. Over the last

decade, NASA study teams have assessed a

variety of mission and technology options for

human exploration missions to the Moon and Mars.

In FY1988, NASA's Office of Exploration sponsored

four separate Exploration Case Studies 2,3 which

outlined strategies for human expeditions to Phobos

and Mars, a human-tended lunar observatory, and

an evolutionary expansion strategy beginning with a

lunar outpost and progressing to similar bases of

operations on Mars and its moons. Phobos mission

objectives included basic exploration, resource surveys to determine the existence of water, and the

establishment of a science station. For the Mars /

Phobos missions, a "split / sprint" transportation

approach was utilized that predeployed cargo

using "minimum-energy" trajectories to reduce

propellant mass, and higher energy trajectories to

reduce in-space transit times for the crew. Short

stay time, opposition-class missions employing

aerobraking, chemical and NTR propulsion

options were also assumed.

The Exploration Case Studies were followed in

1989 by NASA's "90-Day" Study, 4 which focussed

primarily on the establishment of a permanent

lunar base and "all-up" exploration missions

to Mars. "All-up" refers to an operational mode in

which all of the payload and propellant required

for the entire Mars mission is carried on a

single vehicle. The expendable chemical / aerobrake option used direct capsule reentry at Earth

for crew recovery and had an initial mass in low

Earth orbit (IMLEO) of ~831 t. The chemical TMI

stage utilized LOX/LH2 propulsion, and two large

diameter (~ 30 m) aerobrakes, constructed in low

Earth orbit, were used to capture the piloted

lander / ascent vehicle and LOX / LH2 trans-Earth

injection (TEl) stage into Mars orbit. The "all NTR"

option 5 used a single 75 klbf engine for all primary

propulsion maneuvers, including Earth orbit

capture (EOC), and had an IMLEO of ~668 t.

In May 1991, the Synthesis Group issued its

report 6entitled "America at the Threshold: America's

Space Exploration Initiative." In it different architectural approaches and technical strategies were

outlined and fourteen key technologies necessary

for safe and cost effective exploration of the Moon

and Mars were identified. The top two technologies

listed were a heavy lift launch vehicle and NTR

propulsion. The Synthesis report stated that for

Mars transit "the nuclear thermal rocket is the preferred propulsion system allowing significantly

reduced mass to low Earth orbit, shorter transit

times and greater operational flexibility. ''6 The use of

aerobraking for Mars orbit capture (MOC) was

rejected by the Synthesis Group in favor of capture

using NTR propulsion because of a variety of mission-, spacecraft design-, and safety-related issues

associated with aerobraking. 6

In FY93, an intercenter NASA Mars Study Team

was organized by the Exploration Project Office

(ExPO) at the Johnson Space Center (JSC) and

tasked with assessing the requirements for a

piloted mission to Mars as early as 2010. A split /

sprint mission with predeployed cargo was

baselined and NTR propulsion was selected for

all primary propulsion maneuvers in keeping with

the Synthesis Group recommendations. "Fast

conjunction-class" trajectories 7,8 were also featured to maximize the exploration time at Mars

while reducing the total "in-space" transit time to

approximately one year.

The reference Mars architecture was later

changed by ExPO to incorporate a common,

"dual use" aerobrake / descent shell and "in-situ"

resource utilization (ISRU) in an effort to achieve a

single launch cargo and piloted mission capability

using a 240 t-class heavy lift launch vehicle

(HLLV). Common habitat modules were also

assumed for the piloted lander, surface hab and

ERV. Using LH2 brought from Earth, an ISRU plant

would convert Martian carbon dioxide into liquid

oxygen / methane (LOX/CH4) propellant to fuel a

"dry" ascent stage carried to the Mars surface on

the cargo lander mission 9. A second cargo lander

provided an additional habitat module, science

equipment and consumables needed to support the

crew during the long (~500 day) Mars surface

exploration phase. A separate ERV, placed in

Mars orbit, returned the crew and "dual use"

ascent stage crew capsule to Earth where it

provided a direct Earth entry capability. LOX/CH4

propulsion was used on both the descent and TEl

stages to maximize hardware commonality, and

NASA/TM-- 1998- 208834/REV 1 2
NTR propulsion was used only for the TMI stage.

Additional details on the FY93 reference Mars

architecture are provided elsewhere. 1°,11

A common TMI stage powered by three to four

15 klbf NTR engines was developed for both the

cargo and piloted missions 1° (see Figure 1). The

TMI stage was sized by the 2009 piloted mission

and its more energetically demanding 180-day

trajectory and then used in the minimum energy

cargo missions to maximize payload delivery to

Mars. After a "2-perigee burn" Earth departure,

the spent TMI stage was jettisoned and targeted

for long-duration disposal into heliocentric space.

In addition to the reference Mars architecture,

GRC developed "all NTR" mission options (to

capitalize on the NTR's higher performance) and

modular vehicle designs using "standardized"

engine and stage components. 1° The "modular

approach" provided a number of attractive

features which included enhanced mission flexibility

and safety, simplified vehicle design and

assembly, and reduced development / procurement costs through standardization of the "fewest

number" of components. Vehicle designs compatible with a 120 t-class HLLV were also developed

and utilized a dual launch, Earth orbit rendezvous

and dock (EOR&D) scenario for vehicle assembly.

Particularly noteworthy, was the introduction and

integration of "bimodal" NTR engines and active LH2

refrigeration systems into the basic design of the

ERV 1° (see Figure 2). The elimination of boiloff over

the ~4.1 year mission duration of the ERV led to

dramatic reductions in IMLEO, total engine burn

time and LH2 tank size.

In FY97, NASA's intercenter Mars Human

Exploration Study Team was reconvened to

reevaluate, refine and update the FY93 DRM.

Key mission changes 12 included the use of an

~80 t -class HLLV called "Magnum" and adoption

of a dual launch EOR&D vehicle assembly

scenario. Payload manifests, including crew

accommodations and consumables, were critically

examined on each cargo and piloted mission to

save mass and eliminate duplications. Mass

reductions in large structures, like propellant

tanks and habitat modules, were achieved

through the use of advanced composites. A

lightweight, inflatable hab module design

developed by JSC was also examined.

The expendable NTR TMI stage and "new" bimodal

2007 Cargo Mission 1

"Dry" Ascent

Stage and Lander

15.0 m

86.0 t LH2

20.6 m* (at 100%)

--10 m

I[III [I [ H[IIIII_|

4.7 m

__t.---

IMLEO -- 216.6 t

2007 Cargo Mission 2

Hab Module and

Lander

16.3 m

86.0 t LH2

(at 100o/0) I

_10 m --,J

216.6 t

-V

19.0 m

2007 Cargo Mission 3

LOX/CH 4 TEIS and

Hab

86.0 t LH2

{at 91.9%)

_10 m--,-

204.7 t

l

12.0m

1

2009 Piloted Mission 1

Piloted MEV and

Surface Hab

. --.

86.0 t LH2_

(at 100%)

_10m_

[iHIII I[I i I I I ( [ I[lllll

212.1 t

*Expendable TMI Stage LH2 Tank (at 18.2 m length) sized by 2009 Mars Piloted Mission.

Figure 1.--Reference Mars Cargo and Piloted Vehicles Using Common "NTR-Powered" TMI Stage.

NASA/TM-- 1998-208834/REV 1 3
Figure 2.--Artist's Illustration of ERV with 50 kWe "Bimodal" NTR System and Active LH2

Refrigeration. A 5 kWe Solar Array is Shown on the ERV for Scale.

NTR vehicle concepts developed during this

study period were sized to fit within the mass and

payload volume limits of the Magnum HLLV. To

circumvent volume limitations, "LOX-augmented"

NTR (LANTR) engines were also examined to

increase "bulk" propellant density and maximize

vehicle performance while staying within the mass

limitations of a two Magnum scenario.

This paper describes the NTR vehicle and

mission analysis results performed by the Glenn

Research Center over the last ~18 months in

support of NASA's intercenter Mars study effort.

The paper first describes the operating principles

and characteristics of the small, 15 klbf solid core

NTR engines baselined in the study. This is

followed by a discussion of the operational

characteristics and benefits of the "bimodar' NTR

and LANTR engine concepts. Next, key features

of the Mars DRM are reviewed and a summary of

mission and transportation system ground rules

and assumptions are provided. Representative

vehicle concepts and their operational characteristics are then presented for an expendable

NTR TMI stage, several bimodal NTR vehicle

options, and a LANTR vehicle configuration

capable of adding Phobos rendezvous and

landing options to the current DRM. The paper

concludes with a summary of our findings and a

brief discussion of the evolvability of bimodal

LANTR vehicles to support a fully reusable, Mars

mission architecture and future human expansion.

NUCLEAR THERMAL ROCKET PROPULSION

The "solid core" NTR represents the next major

evolutionary step in propulsion technology and is

key to providing "low cost access through space"

for future human exploration missions to the

Moon, Near Earth Asteroids and Mars. The NTR

is not a new technology. Its feasibility was

convincingly demonstrated in the United States

during the Rover / NERVA (Nuclear Engine for

Rocket Vehicle Application) nuclear rocket

programs. 13 From 1955 until the program was

stopped in 1973, a total of twenty rocket reactors

were designed, built and tested. These integrated

reactor / engine tests, using LH2 as both reactor

coolant and propellant, demonstrated a wide range

of engine sizes (from ~50 to 250 klbf), high

temperature graphite fuel providing substantial

hydrogen exhaust temperatures (~2350-2550 K),

NASA/TM-- 1998- 208834/REV 1 4
sustained engine operation (over 60 minutes for a

single burn) and restart capability (28 startups and

shutdowns on the NRX-XE engine). The Rover /

NERVA program costs were estimated at

~$1.4 billion (an ~$10 billion investment today).

Approximately four years after the start of the

NERVA program, a nuclear rocket program was

initiated in the former Soviet Union known today

as the Commonwealth of Independent States

(CIS). 14 Extensive nuclear and non-nuclear

subsystem tests were conducted, including fuel

element and reactor tests at the Semipalatinsk

facility in Kazakhstan. 15 Although no integrated

engine system tests were conducted, a high

temperature ternary carbide fuel element was

developed capable of producing hydrogen

exhaust temperatures in excess of 3000 K-- about

500 K higher than the best NERVA fuels.

NTR Operating Principles

Conceptually, the NTR engine is relatively

simple (see Figure 3). High pressure propellant

flowing from pumps cools the nozzle, reactor

pressure vessel, neutron reflector, control drums,

core support structure and internal radiation

shield, and in the process picks up heat to drive

the turbines. The hydrogen exhaust is then

routed through coolant channels in the reactor

core's fuel elements where it absorbs the energy

released by fissioning uranium atoms, is

superheated (to 2700-3100 K), and then

expanded out a supersonic nozzle for thrust.

Controlling the NTR engine during its operational

phases (startup, full thrust, and shutdown) is

accomplished by matching the turbopumpsupplied hydrogen flow to the reactor power level.

Control drums, located in the surrounding

reflector region, regulate the number of fissionreleased neutrons that are reflected back into the

core and hence the reactor power level. An

internal neutron and gamma radiation shield,

containing interior coolant passages, is also

placed between the reactor core and sensitive

engine components to prevent excessive

radiation heating and material damage.

Ternary Carbide Fuel NTR Engine Design

What's new about NTR propulsion today that

warrants renewed investment in this technology?

The answer lies in a reduced size, higher

performance engine that can be ground tested at

full power in a "contained facility" meeting current

environmental regulations. Design studies, 16,17

funded by NASA's Nuclear Propulsion Office in

1992-1993 and conducted by a US / CIS industry

team of Aerojet, Energopool and Babcock and

Wilcox (B&W), produced a small advanced NTR

engine concept with impressive parameters:

thrust ~15 klbf, Isp ~940-960 s, engine thrust-toweight ~3.1, and "full power" engine fuel lifetime

of ~4.5 hours. The CIS engine design (shown in

Reactor--_ _ / Radiatio_Shield

t -

__N _---__--___-_.: "l_ LaH_:

Reflector--_ Control Drum _ Pumps -_

-Turbines

Figure 3.--Schematic of "Solid Core" NTR Using Dual Turbopump Expander Cycle.

NASA/TM-- 1998-208834/REV 1 5
From TPA Pump

2

3

4

6

8

.... ili

TPA

Turbine Turbine

Figure 4.--Component Layout / Flow Schematic

1 - Nozzle

2 - Closing Device

3 - Core Support Structure

4-FA

5 - Injector

6 - Moderator

7 - Pressure Vessel

8 - Control Drum

9 - Side Reflector

10 - End Reflector

11 - Radiation Shield

12 - Safety Rod

13 - Recuperator

of CIS Engine.

Figure 4) utilizes a heterogeneous reactor core

design with hydrogen-cooled zirconium hydride

(ZrH) moderator and ternary carbide fuel materials.

The ZrH moderator is located between reactor fuel

assemblies and is very efficient in minimizing the

inventory of fissile material in the reactor core. The

CIS fuel assembly is an axial flow design and

contains a series of stacked 45 mm diameter

bundles of thin (~1 mm) "twisted ribbon" fuel

elements approximately 2 mm in width by 100 mm

in length. The "fueled length" and power output

from each assembly is determined by specifying

the engine thrust level and hydrogen exhaust

temperature (or desired Isp). For a 15 klbf engine,

36 fuel assemblies (with 6 fuel bundles each) are

used to generate the required 335 MWt of reactor

power at the same Isp.

The ternary or "tricarbide" fuel material in each

"twisted ribbon" element is composed of a solid

solution of uranium, zirconium and niobium

carbides having a maximum operating temperature expected to be about 3200 K. The fuel

composition along the fuel assembly length is

tailored to provide increased power generation

where the propellant temperature is low, and

reduced power output near the bottom of the fuel

assembly where the propellant is nearing its

exhaust temperature design limit. In this current

study, the CIS engine total power output has

been fixed at 335 MWt and the hydrogen exhaust

temperature allowed to vary from 2900 to 3075 K

to provided increased Isp operation (from ~940

to 955 s) when needed. During reactor tests,

hydrogen exhaust temperatures of 3100 K for

over one hour and 2000 K for 2000 hours were

demonstrated in the CIS. 14

CIS Enqine Power Cycle / Desiqn Characteristics

The CIS engine design utilizes a dual

turbopump, "recuperated" topping cycle. 16,17

Hydrogen flowing from each pump is split (see

Figure 5), with ~84% of the flow going to a

combination recuperator/gamma radiation shield

and the remaining 16% used to cool the nozzle.

The recuperator / shield, located at the top of the

engine, provides all of the necessary turbine drive

NASA/TM-- 1998- 208834/REV 1 6
772 R

2954 psia

1155 R 1155 R

2719 psia 2719 psia

81.5% Flow 18.5% Flow

Qmod (5.3°/o)_ 3 Qref (1•2°/o_'_l

_ Moderator I I Reflector l

l 779R 2954 psia t

Qpv (0.3 _..%)_%_,782 R Qins |2954 psia (0.3%)

I Press-.Vessel I

$ 760 a _

. |3042 psia t Internal

[ Turbine J INeutron She d

844 R

"_'_ 4800 psia

Qnoz (2°_ 84% Flow

I NozzlT---1

80 R

4947 psia

16% Flow

32.4 R

32.3 psia

..oo.0°°°°°0°ooo00...,,....°.._

Qrec (11.9%)

Qfa _=

(92.2%)_._

1155 R

2719 psia

100% Flow

499 R

' 2542 psia

100% Flow 5220 R

Pc = 2000 psia

Figure 5.--Flow Schematic of Recuperated Topping Cycle for the CIS

Engine.

power. The turbine exhaust cools the reactor

pressure vessel and is then merged with the

nozzle coolant to cool the moderator and reflector

regions of the engine. The coolant then passes

through borated ZrH and lithium hydride (LiH)

neutron shields located within the pressure

vessel between the reactor core and the

recuperator/gamma shield (see Figures 4 and 5),

before returning to the recuperator where it heats

the pump discharge flow. Exiting the recuperator,

the cooled hydrogen is then routed to the core

fuel assemblies where it is heated to the required

design temperatures. The 15 klbf CIS engine

design has a chamber pressure of 2000 psia, a

nozzle area ratio of 300 to 1, and a 110% bell

length nozzle resulting in Isp values of ~940 to

955 s for hydrogen exhaust temperatures in the

range of 2900-3075 K. The approximate engine

length and nozzle exit diameter for the 15 klbf CIS

engine is ~4.3 m and ~1.0 m, respectively.

A summary of key design features of the CIS engine

is found in Table 1.

The "Bimodar' NTR--A Fully Inteqrated System

The bimodal NTR engine and vehicle concept

was examined in detail during this study period

to more fully exploit the performance potential

of the NTR and enhance stage capabilities.

Besides its impressive propulsion characteristics,

the solid core NTR represents a "rich source of

energy" because it contains substantially more

uranium-235 fuel in its reactor core than it

consumes during its primary propulsion

maneuvers. By reconfiguring the NTR engine for

"bimodar' operation (Figure 6), abundant electrical

power can also be generated to support spacecraft environmental systems, high data rate

communications, and enhanced stage operations

such as an active refrigeration / reliquification

system for long term, "zero-boiloff" LH2 storage.

A bimodal NTR-powered spacecraft would be very

similar to today's nuclear-powered submarine

which uses high-pressure steam provided to a

turbine engine to drive the submarine's propeller.

Steam from the reactor also generates all of the

submarine's electricity.

Besides providing a continuous source of

reactor thermal energy, bimodal operation is also

beneficial because it: 1) reduces thermal stress on

the reactor (it's pre-heated); 2) minimizes large

thermal cycling (no prolonged, deep "cold soak"

of the engine); 3) allows rapid reactor restart (in

case of emergency); 4) minimizes "decay heat

removal" propellant penalty (by rejecting low

power, "after-heat" through the power system's

space radiator); and 5) provides a source of

NASA/TM-- 1998-208834/REV 1 7
Table 1. Key Design Features of CIS / NTR Engines

Reactor Power

Engine Thrust (klbf)

Hydrogen Exhaust Temperature, K

Propellant Flow Rate, kg/s

Specific Impulse, s

Fuel Composition

Fuel Form ("Twisted Ribbon"), mm

Fuel Element Power Density (ave), MW/L

Core Power Density, MW/L

Fuel Volume, liters

Number of Assemblies (Elements)

Number of Safety Rods

Vessel Diameter, m

Reactor Fueled Length, cm

Reactor Mass (with internal

shielding and recuperator), kg

Engine Thrust-to-Weight Ratio

Total Engine Length, m

Nozzle Exit Diameter, m

335

(15-14.76)

2,900 - 3,075

7.24 - 7.01

940 - 955

(U,Nb,Zr)C

Approximate 100 x 1.6 x 1.0

3O

5.0

11.5

36

13

0.65

55

2224

3.06

4.3

1.0

heated, gaseous hydrogen (GH2) for propellant

tank pressurization, and possible high Isp attitude

control and orbital maneuvering systems.

During the power generation phase, the

bimodal engine's reactor core operates in

essentially an "idle mode" with a thermal power

output of ~110 kilowatts. The energy generated

within the reactor fuel assemblies would be

removed using a variety of "closed loop" concept

options (such as core support tie tubes,

integrated energy extraction ducts within the

individual fuel assemblies, or a throat closure plug)

and then routed to a turboalternator-compressor

Brayton power conversion unit using a heliumxenon (He-Xe), hydrogen-nitrogen (H2-N2), or

other working fluid combination (see in Figure 6).

A pumped-loop radiator system is used to reject

system waste heat and is also available to help

remove low level decay heat power following

high thrust engine operation.

Several options for closed Brayton cycle (CBC)

power generation are being considered for the

CIS engine design. Although the current CIS/

CBC system is designed to radiate small amounts

of thermal power at lower temperature (~1300 K)

during the electric power generation phase, the

same system can reject several megawatts of

decay heat by operating the radiators at higher

temperatures since heat transfer to space depends on the radiator surface temperature raised

to the fourth power. Molybdenum alloy turbine

wheels and niobium alloy static structures can

withstand 1400 to 1500 K GH2 inlet temperatures

because the materials are compatible with GH2

and have high strength-to-density ratios at these

temperatures. 16,17 Within an hour or two after

thrust generation, reactor power decays significantly and the CIS / CBC temperatures drop. For

decay heat removal or higher power mode

operation, coolant is routed through the fuel

assemblies (FA) after the CIS Brayton cycle loop is

closed by inserting a nozzle "throat plug" located

at the aft end of a central drive shaft (see

Figure 7). This action opens an annular duct which

carries the coolant / working fluid to the CBC

turbine inlet. 16,17 In order to prevent excessive

NASA/TM-- 1998- 208834/REV 1 8
_ Refrigeration

Generator

m _ Power

Conditioning

Compressor H2 Turbopump I I__ On-board Systems

I

I L_ Payload

Im m .[] Electric Propulsion

Radiator Heat Turbine ,_ used on Hybrid

Exchanger BNTR/EP option

Brayton Power Conversion is

Primary Option/Other Options

Can Also Be Adapted 4

Thermal

Propulsion

4

Electric

Propulsion

Figure 6.--Key Components of "Bimodal" NTR Stage -- "A Fully Integrated System."

From Brayton Cycle

1

2

3

4

6

8

12

! Brayton

To Brayton Cycle

Cycle

1 - Nozzle

2 - Closing Device

1 3 - Core Support Structure

4-FA

5 - Injector

6 - Moderator

7 - Pressure Vessel

8 - Control Drum

9 - Side Reflector

10 - End Reflector

11 - Radiation Shield

12 - Safety Rod

13 - Recuperator

Figure 7.--Design Features of "Bimodal" CIS Engine Concept.

NASA/TM-- 1998-208834/REV 1 9
loss of coolant past the throat plug during many

months of low electrical power generation, the

GH2 coolant /working fluid is rerouted to passages

through the FA walls before entering the Brayton

rotating unit. During this period, the throat plug

remains closed as a reliability enhancement

feature, inhibiting possible coolant leakage from

the system through any cracks that may develop

in the FA wall.

The "LOX-Auqmented" NTR (LANTR) Concept

An innovative "trimodar' NTR concept, 18,19

known as LANTR, is presently under study by

NASA GRC which combines conventional LH2-

cooled NTR, Brayton cycle power generation and

supersonic combustion ramjet (scramjet) technologies. During LANTR operation, oxygen is

injected into the large divergent section of the

NTR nozzle which functions as an "afterburner"

(see Figure 8). Here, it burns spontaneously with

the reactor-heated hydrogen emerging from the

LANTR's sonic throat adding both mass and

chemical energy to the rocket exhaust--essentially

"scramjet propulsion in reverse."

The trimodal LANTR engine, illustrated in

Figure 9, can operate as a conventional

LH2-cooled NTR, a bipropellant LOX/LH2 engine

and a power reactor. Its prinicipal components include a reactor and nozzle to heat and expand

propellant, hydrogen and oxygen tankage and feed

systems (using autogenous gas bleed for tank

pressurization), and a closed Brayton cycle system

for electric power generation and deep throttling.

The CBC can also be used for engine "cooldown"

assist as discussed above. The hydrogen feed

system is powered by engine waste heat using the

CIS recuperated topping cycle which enables the

engine to run at a nozzle inlet pressure of 2000 psia.

This and the fact that the recuperator also

doubles as the reactor's cooled gamma radiation

shield helps reduce engine size and mass. The

LANTR engine generates electricity by bleeding

reactor-heated GH2 or other working fluid through

the Brayton cycle turbine, which drives an electric

motor / generator and compressor. An "on-off"

valve or throat plug is used to shut the nozzle

throat during CBC operation and prevent leakage

of the working fluid to space, and opened to

the hot hydrogen exhaust during thrust mode

Reactor .--_ Nozzle

Core _ Throat\

_ _ LOX _ _SuPers°Nc

_ _,,\ _ _v_.,_-_,_ Combustion

"____ ._--I-_ and

Hydrogen _ _ _ --Supersonic_..._ " _ _._,,..u

Coolant/ _ _ _ --'-.'. ...... .,,--.6--_ Thrust

/ J" 1-12 Augmentation Propellant _ _ /.__ ._ Augm_entati_on_._

----Z j In,LOtXon__._ ATc ->3'500 K)

Subsonic

Hot H2

(Tex~ 2,600 to 3,100 K)

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*For 15 klbf LANTR with chamber pressure = 2,000 psia and _ = 500 to

Figure 8.--Schematic/Characteristics of "LOX-Augmented" NTR.

NASA/TM-- 1998- 208834/REV 1 10
Nozzle

Radiator

A Q-----_

Heat Pipe

Radiator

r-_

"1

Electri

r

Turbo-

]Compressor

31 Motor-

.) Generator

H2

TPA

I

Igniter

0 2 Rich

Preburner

1

Reactor Fuel

Assemblies

0 2 Injector

LO 2

02

TPA

Brayton

Power System

Reactor and

H 2 Feed System Nozzle

H2 Feed System

Brayton Power System

0 2 Feed System

.......... Tank Pressurization

Figure 9.--Flow Schematic of "Bimodal" LANTR Engine.

operation. Waste heat can be rejected to space

using a combination of nozzle and heat pipe

radiator (as shown in Figure 9), or a dedicated

radiator system as assumed in this study.

During bipropellant operation the oxygen feed

system uses a topping cycle powered by an

oxidizer-rich preburner. Downstream nozzle

injection isolates the reactor core from oxygen

damage provided the throat retains choked flow.

This condition is satisfied by using a "cascade"

scramjet injector concept developed by Aerojet

which controls oxygen addition and heat release

profiles (via staged injection) to keep the flow

supersonic. 18 It also increases penetration, mixing

and combustion of the oxygen injectant in the

supersonic hydrogen flow while minimizing shock

losses and formation of high heat flux regions (hot

spots), thereby maximizing engine performance

and life. The high reactor outlet pressure of the

LANTR (~2000 psia) also enables high area ratio

nozzles (_ = 500 to 1), important for combustion

efficiency, at reasonable size and mass.

The LANTR concept has the potential to be an

extremely versatile propulsion system. By varying

the engine's oxygen-to-hydrogen (O/H) mixture

ratio (MR), LANTR can operate overa wide range of

thrust and Isp values (Figure 8) while the reactor

core produces a relatively constant power output.

For example, as the MR varies from 0 to 7, the

engine thrust-to-weight ratio for a 15 klbf NTR

increases by ~440%--from 3 to 13--while the Isp

decreases by only ~45%--from 940 to 515 seconds.

This thrust augmentation feature means

that "big engine" performance can be obtained

using smaller, more affordable, LH2-cooled NTR

engines that are easier to develop and test in

"contained" ground facilities. The engines can

then be operated in space in the augmented high

thrust mode to shorten burn times (thereby

extending engine life) and reduce gravity losses

(thereby eliminating the need for and concern

over multiple, "perigee burn" Earth departure

maneuvers). Reactor preheating of hydrogen

before oxygen injection and combustion also

results in higher Isp values than found in LOX / LH2

NASA/TM-- 1998-208834/REV 1 11
chemical engines operating at the same mixture

ratio (~100 s at MR = 6). Lastly, the ability to

substitute high-density LOX for low-density LH2

provides the vehicle designer substantial flexibility

in configuring spacecraft which can accommodate

a wide variety of mission needs, as well as,

"volume-constrained" launch vehicle designs.

DESIGN REFERENCE MISSION DESCRIPTION

The Mars Exploration Study Team is presently

assessing a variety of mission architectures and

transportation system options for conducting a

human mission to Mars in the 2014 timeframe

centered around a split cargo/piloted sprint mission

approach. The mission profile shown in Figure 10

assumes the use of aerobraking at Mars and "insitu" production of ascent propellants to reduce

mission mass and transportation system requirements from Earth. The piloted mission is preceded

by two cargo missions which depart Earth in

November 2011 and arrive at Mars ~297 days later.

Each cargo and piloted vehicle requires two ~80 t

"Magnum" HLLVs (one for the aerobraked payload

and the other for the NTR TMI stage) and utilizes an

EOR&D vehicle assembly sequence. A "common"

aerobrake / descent shell is assumed for either

capture into Mars orbit or direct descent to the

Mars surface. The expendable NTR TMI stage (not

shown in Figure 10) is jettisoned after an

appropriate "cooldown" period and subsequently

disposed of along its heliocentric trajectory.

The cargo lander mission carries a surface

payload consisting of a "dry" Mars ascent stage and

crew cab combination, nuclear power systems, LH2

"feedstock" and ISRU plant, an inflatable laboratory

module, rovers and science equipment (The

complete mass manifest for the cargo lander is

found in the Appendix in Table A-2). The payload

element delivered to Mars orbit consists of the

crew return habitat module, "fueled" TEl stage

Cargo delivered to LEO

on Magnum, rendezvous

with NTR.

Outbound Hab delivered to LEO on

Magnum. Crew of 6 delivered to LEO

in shuttle or other. Both rendezvous

with NTR.

.... • ,.,,,,..

"-.

Grew direct enters in

capsule, Apollo-style.

Earth return Hab is

discarded

Opportunity 1 (2011): 2 flights ,_m_ v_':_'argo, 2

Return habitat, chemical _.-_ --_-'_p" ."V

TEl stage, (aerocapture /' - ""f'_'"ars

to 1 Sol orbit) ." /

Ascent production, vehiclesurface prop !.:: [ _/_Cargo 1

exploration gear, i

inflatable Hab and "

surface nuclear power

(direct entry to surface)

Opportunity 2 (2014): 1 flight

Crew of 6 aerocaptures into 1 Sol

orbit, then lands in outbound Hab.

Surface rendezvous with

pre-deployed assets.

_.Crew ascends

to return Hab

.,_ in capsule

Figure 10.--Candidate Mission Profile for Mars Design Reference Mission.

NASA/TM-- 1998- 208834/REV 1 12
and integrated aerobrake structure. After the

operational functions of the ERV and cargo lander

are verified, and the ascent stage is fully fueled

with LOX/CH4 propellant, the piloted vehicle leaves

Earth in January 2014 (mass manifests for the ERV

and piloted lander are found in the Appendix

Tables A-1 and A-3, respectively). It arrives at Mars

~180 days later using a "fast conjunction-class"

trajectory, 7,8 which maximizes the exploration time

at Mars while reducing the total in-space transit

time to approximately a year. After a 554-day stay at

Mars, the crew returns in the ascent portion of the

cargo lander to a waiting ERV to begin preparations

for the 6 month journey back to Earth. The ascent

stage crew cab doubles as an Earth crew return

vehicle (ECRV) and is retained by the ERV for the

trip home. Nearing Earth, the crew separates from

the ERV and reenters the atmosphere in the

ECRV while the ERV flys by Earth and continues

on into deep space. The total duration of the

piloted and ERV missions are 914 days and

1701 days, respectively.

MARS MISSION / TRANSPORTATION SYSTEM

GROUND RULES AND ASSUMPTIONS

The ground rules and assumptions for the

reference mission architecture and NTR-based

transportation system examined in this study are

summarized in Tables 2 and 3, respectively. In

Table 4, the AV budgets are listed for both the

aerobrake (AB) and "propulsive capture" (PC)

versions of the DRM. Table 5 provides additional

AV requirements for the "all NTR" mission options

which take into account disposal of spent cargo

and piloted NTR stages (either along their interplanetary trajectories or into a stable heliocentric

orbit between Earth and Mars at 1.19 astronomical

units [A.U.]) at mission end. While Table 2 highlights key features and characteristics of the DRM

(e.g., scaling of the" triconic" aerobrake/descent

shell mass), Table 3 provides details on NTR

and LANTR systems, auxiliary RCS propulsion,

cryogenic tankage, propellant thermal protection

and boiloff rates, refrigeration system mass and

power requirements, and contingency factors

used in this study. Although primary propulsion

maneuvers are performed using either the NTR or

LANTR engines, the spacecraft also executes

midcourse and secondary maneuvers using a

storable, bipropellant RCS system.

The use of composite materials is assumed for all

Mars transportation stage masses (e.g., descent /

ascent stages, NTR LH2 propellant tanks and

primary structures, etc.) for weight reduction. The

wall thicknesses for the LH2 tanks were calculated

based on a 35 psi internal pressure and included

hydrostatic loads using a "5g" loading and a safety

factor of 1.5. A 3 percent ullage factor was also

baselined in this study. For the LOX tanks on

LANTR, a 50 psi internal pressure was assumed

resulting in wall thicknesses of ~0.05 inches.

An 80 layer (~2.1 inch), multilayer insulation

(MLI) system (at 38 layers per inch) is assumed for

thermal protection 2° of the LH2 and LOX cryogenic

tanks. This insulation thickness exceeds the

"ground hold" thermal protection requirements for

"wet-launched" LH2 tanks which need a minimum

of ~1.5 inches of helium-purged insulation. 21 The

installed density of the 80 layer MLI system is

~1.44 kg / m 2, and the resulting LH2 boiloff rate in

LEO is ~3.11 x 10-2 kg/m2/day (based on an

estimated heat flux of ~0.161 W / m 2 at a LEO sink

temperature of ~230 K). The corresponding boiloff

rate for LOX is shown in Table 3. Finally, to account

for micrometeoroid protection of propellant tanks

(while in LEO, Mars orbit, and during transit to and

from Mars), an ~0.50 mm thick sheet of aluminum

(corresponding mass of ~1.35 kg / m 2) is also

included in the total tank weight estimates.

The NTR vehicle concepts developed in this

study employ different thermal protection systems

for LH2 consistent with the vehicle's mission

application and expected lifetime. For the expendable NTR TMI stages, which have a "limited life" in

LEO of ~32 days before departure, an ~2 inch

"minimum mass" MLI system is used resulting in

a LH2 boiloff of ~0.46 t. The "all BNTR"-powered

ERV mission has the most demanding requirements for thermal protection with a mission

ellapsed time between TMI and TEl of 1521 days

(~4.2 years). For this mission application, an active

system was developed consisting of a 2 inch MLI

blanket and a turbo-Brayton refrigerator. Selection

of the turbo-Brayton system was based on a

NASA-funded study and survey 22 of various

refrigeration systems which indicated its suitability

for large LH2 tanks requiring refrigeration capacities

in the 10 to 100 watt cooling range. Table 3 shows

the specific mass and input power assumptions

used in estimating the inert weight and electrical

power demands for the common, "refrigerated"

BNTR core stage developed in this study.

NASA/TM-- 1998-208834/REV 1 13
Table 2. Mars Mission Study Ground Rules and Assumptions

Split Mission Scenario: (2 Cargo Missions in 2011,

1 Piloted Mission in 2014)

Payload Elements Consist of Mars Cargo Lander, Earth Return Vehicle (ERV)

and Piloted Lander with 6 Crew.

Dual Launch Earth Orbit Rendezvous and Dock Vehicle Assembly at 407 km

using two 80 - 88 t "Magnum" HLLVs.

Magnum Payload Shroud Dimensions:

7.6 m (I.D.) x ~ 28.0 m Length

Aerobraking and Propulsive Capture into 250 x 33,793 km (1 sol) Elliptical

Mars Parking Orbit

Aerodescent Shell and Parachutes for Descent to Mars (descent AV = 632 m/s)

Aerobrake/Descent Shell Sizing: MAB(t)= _/MpL (a + bYe) + Ms; where

(MpL= payload mass in t, a = -0.55, b = 0.19, Ve = entry velocity in km/s and

Ms = structural mass = 6 t)

Mars Descent Stage uses 4 - 15 klbf LOX/CH 4 Engines (Isp = 379 s, MR = 3.5,

Stage Boiloff Rate: ~ 0.4 %/month)

"ln-Situ" Production of LOX/CH4 Ascent Propellant using Earth-Supplied LH2

Mars Ascent Stage AV to 1 sol orbit: 5625 m/s

Mars Ascent Stage AV to Phobos orbit: 5400 m/s

Mars Ascent Stage and Crew Capsule Rendezvous with ERV/Crew Capsule

Retained/Doubles as Earth Crew Return Vehicle (ECRV)

Chemical Trans-Earth Injection (TEl) Stage uses 2 - 15 klbf LOX/CH 4 Engines

(Stage Boiloff Rate: ~ 0.2%/month)

Direct Reentry of ECRV and Crew at Earth Arrival

Mission Abort Strategy:

Outbound: Abort to Mars Surface

At Mars: Abort to ERV, which carries contingency consumables.

NASA/TM-- 1998- 208834/REV 1 14
Table 3. Mars NTR / LANTR Transportation System Assumptions

NTR / LANTR Thrust/Weight = 15 klbf / 2224 kg (LH2 NTR)

Systems: = 15 klbf/2630 kg (LANTR @ MR = 0.0)

Fuel / Propellants = Ternary Carbide / Cryogenic LH2 & LOX

Isp = 940 - 955 s (@ O/F MR = 0.0 / LH2 only)

= 831 s (@ O/F MR = 0.5)

External Shield Mass = 2.84 kg/MWt of reactor power

Flight Reserve = 1% on AV

Residual = 2% of total tank capacity

Cooldown (effective) = 3% of usable LH2 propellant

RCS System: Propellant = a204 / MMH

Isp = 320 s

Tankage = 5% of total RCS propellants

Cryogenic Material = Advanced Composite

Tankage/ Diameter = 7.4 m (LH2) / 2.6 m (LOX)

Thermal Geometry = cylindrical with _/2/2 domes / spherical

Protection: Insulation = 2.1 inches (80 layers) MLI @ 1.44 kg/m 2

LH2/LOX Boiloff* = 3.11 x 10 2/ 6.49 x 10 2 kg/m2/day

Shield = 1.35 kg/m 2 (~0.5 mm sheet of Aluminum)

LH2 Refrigeration Specific Mass = 4.57 kg/W refrig. @ 75 Watts

System: Input Power = ~0.11 - 0.20 kWe / W refrig.

Contingency Engine, shields and stage dry mass = 15%

* Based on estimated heat flux of ~ 0.1608 W/m 2at LEO sink temperature of ~230 K

Table 4. Mars Cargo and Piloted Mission AV Budgets (Ideal)

Outbound Inbound Total TMI MOC TEI/EOC Total Vehicle Launch Transit Time Transit Time Mission Time AV AV AV Ideal AV

Mission Mode Date

(days) (days) (days) (km/s) (km/s) (km/s) (km/s)

Cargo 11/8/11 297 NA 297 3.580 AB NA 3.580

(AB @ Mars)

11/9/11 307 NA 307 3.581 0.925 NA 4.505

(PC @ Mars)

12/4/13 294 NA 294 3.605 1.162 NA 4.767

(PC @ Mars)

12/31/13 328 NA 328 3.572 AB NA 3.572

(AB @ Mars)

Piloted 1/4/14 180 180 914 3.672 AB NA 3.672

(AB @ Mars) (554 @ Mars)

2/2/14 180 180 885 4.214 2.251 NA 6.465

(PC @ Mars) (525 @ Mars)

1/21/14 210 180 897 3.861 1.720 NA 5.581

(PC @ Mars) (507 @ Mars)

1/18/14 220 180 900 3.823 1.629 NA 5.452

(PC @ Mars) (500 @ Mars)

E RV 11/8/11 297 180 1702 3.580 AB 1.079 4.659

Outbound/ (AB @ Mars) (1225 @ Mars)

Piloted 11/9/11 307 180 1701 3.581 0.925 1.079 5.585

Inbound (PC @ Mars) (1214 @ Mars)

11/9/11 307 180 1731 3.581 0.925 1.419/1.365 7.290

(PC @ Mars) (1244 @ Mars)

Note:

AV based on 407 km circular orbit at Earth and 250 X 33793 km Mars parking orbit.

G-losses appropriate to "single or double perigee burn" Earth departure must be added to the TMI AV shown.

Apsidal/nodal alignment penalty of 500 m/s must be added to the TEl AV value shown.

NASA/TM-- 1998-208834/REV 1 15
Because of the inventory of radioactive fission

products that will be generated in the BNTR

engines during their service life, care must be

taken to dispose of these vehicles in a responsible manner at mission end. Calculations by

Stancati 23,24 using the Planetary Encounter

Probability Analysis (PEPA) code have provided

estimates of the AV requirements and probabilities

of NTR vehicle collisions with Earth for various

disposal scenarios (shown in Table 5). In the Mars

mission scenario depicted in Figure 10, the

expendable NTR TMI stages are disposed of along

their interplanetary path after payload separation.

Table 5 shows that the probabilities for Earth

reencounter over the course of a million years are

~13% and 11% for the cargo and piloted TMI

stages, respectively. The increased probability for

the cargo missions are due to their near-Hohmann

trajectories. For the "all NTR" mission scenarios,

the BNTR stages used on cargo and piloted lander

missions are removed from Mars orbit shortly after

the ERV leaves for Earth. Although a stable

parking orbit exists at ~1.19 A. U., the AV penalty

for disposal to this location is appreciable at

~2.52 km/s (see Table 5). A second disposal option

adopted in this study is to leave the NTR vehicles

on their flight paths to 1.19 A. U., but to eliminate

the final capture and circularization burns. This

option reduces the disposal AV to ~0.33 km/s and

though it allows for possible future encounters with

Earth, the probabilities are very small (<<1%).

EXPENDABLE TRANS-MARS INJECTION STAGE

A "common" TMI stage design has been

developed for both the Mars cargo and piloted

missions which employs three ~15 klbf CIS / NTR

engines, each weighing 2224 kg and operating

Table 5. Mars Disposal AV Requirements

Mission Disposal Req'd Maneuvers AV Disposal Earth Encounter

Initiated (km/s) Probability

• 2011 Cargo after TMI/ none - TMI stage 0 13% in 106years

(AB @ Mars) before MOC disposed along

interplanetary path

• 2011 Cargo from Mars depart Mars orbit/ 0.331 0

(PC @ Mars) orbit after circularize @ 1.19AU 2.184

cargo delivery 2.bl b

• 2011 Cargo from Mars depart Mars orbit to 0.331 0.02% in 106

(PC @ Mars) orbit after 1.19AU / dispose 0 years

cargo delivery along interplanetary u._l

path

• 2014 Piloted after TMI/ none - TMI stage 0

(AB @ Mars) before MOC disposed along

interplanetary path

11% in 106years

• 2014 Piloted from Mars depart Mars orbit/ 0.331 0

(PC @ Mars) orbit after circularize @ 1.19AU 2.184

cargo delivery 2.bl b

• 2014 Piloted from Mars depart Mars orbit to 0.331 0.02% in 106

(PC @ Mars) orbit after 1.19AU / dispose 0 years

cargo delivery along interplanetary 0.331

path

• 2011 Earth after Earth

Return Stage flyby & ECRV

(PC @ Mars) separation

• 2011 Earth after Earth

Return Stage flyby & ECRV

(PC @ Mars) separation

Earth gravity assist/ 0

circularize @ 1.19AU 2.951

2Lgb1

Earth gravity assist/

disposal along

interplanetary path

0 11% in 106years

NASA/TM-- 1998- 208834/REV 1 16
with an Isp of ~940 s. For a fixed total reactor power

output of ~335 MWt, the engines are capable

of operating at higher Isp values (~955 s) by

increasing fuel temperature (from 2900 K to

~3075 K) which results in a small decrease in thrust

(down to ~14.76 klbf). The single tank stage is

sized to accommodate both the 2007 ERV cargo

mission with a C3 =13.41 km 2 /s 2 and a payload of

~74 t, or the energetically demanding, fast transit

2009 piloted mission (with C3 = 20.06 km 2 / s2).

The size, mass and key features of the common

NTR TMI stage and its aerobraked payloads is

illustrated in Figure 11 and a rendered threedimensional (3-D) image of the stage and payload

is provided in Figure 12. The TMI stage LH2 tank is

cylindrical with _/2/2 ellipsoidal domes. It has an

inner diameter of 7.4 m, an ~20.6 m length, and a

maximum propellant capacity of ~56 t assuming a

3% ullage factor. The main stage components

include the LH2 tank; thermal and micrometeoroid

protection; a forward cylindrical adaptor section

housing avionics and auxiliary power, RCS and

docking systems; forward and aft skirts; thrust

structure; propellant feed system; and NTR

engines. Stage auxiliary power is provided by an

oxygen/hydrogen fuel cell system which supplies

1.5 kWe for up to 32 days in LEO. Assuming a

consumption rate of ~0.415 kg per kWe-hour,

~0.48 t of reactants (at an O / H ratio of 8 to 1) are

required. The hydrogen reactant is drawn from the

main propellant tank while the oxygen reactant is

stored in several small spherical tanks in the

forward section of the stage. The expendable TMI

stage has a length of ~27.5 m as shown in Figure 11

and a total "dry mass" estimated to be ~22.2 t. For

the piloted missions, an external disk shield is

added to each engine to provide crew radiation

protection. This added shielding increases the

stage dry mass by ~3.2 t. A summary mass breakdown for the TMI stage is provided in Table 6.

To minimize LH2 boiloff during the vehicle

assembly phase, the cargo lander and ERV payloads are launched first, followed by the two TMI

stages. Assuming 30 days between Magnum

launches and ~2 days for vehicle checkout, the

longest period any TMI stage is in LEO is ~32 days.

After EOR&D and checkout, the ~51 m long

cargo and piloted vehicles are ready to leave for

Mars. A "2-perigee burn" Earth departure scenario

is assumed which includes gravity losses and a 1%

margin on total TMI AV. The gravity losses for the

cargo lander and ERV missions (C3 ~8.95 km 2/s2),

Table 6. Mass Breakdown for "Common"

NTR TMI Stage*

Stage Element Mass (t)

Structure

Propellant Tank (Lt = 20.6 m x 7.4 m I.D.)

Thermal/Micrometeor Protection System

Avionics and Power

Reaction Control System (RCS)

NTR Assemblies

• 15 klbf CIS NTRs (3)

• External Shields (3)

• Propellant Feed, TVC, etc.

Contingency (15%)

2.45

6.66

1.39

1.2

0.42

6.67

0 - 2.82

0.56

2.90 - 3.33

"Dry" TMI Stage 22.24 - 25.48

LH2Propellant (max. LH2cap. = 56.0 t) 52.0 - 52.61

RCS Propellant 0.77 - 0.88

Fuel Cell Reactants (02) 0.43

"Wet"TMI Stage 75.44 - 79.40

*2007 ERV mission sizes the TMI stage LH2tank.

and the piloted lander mission (C3 ~11.04 km 2 / s2)

are ~95, 110 and 101 m/s, respectively. Similarly,

the corresponding total TMI engine burn times for

the three missions are ~35, 39 and 36 minutes--

well within previously demonstrated capabilities.

Table 7 summarizes the mission mass manifests

for the first two cargo flights and the subsequent

piloted mission. The cargo lander carries the crew

ascent stage (shown in Figure 13) and utilizes a

jettisonable aerobrake / descent shell. It has a total

mass of ~66 t of which ~40.2 t is surface landed

payload. The mass of the aerobrake is estimated to

be ~9.9 t assuming a Mars entry velocity of

~5.65 km/s and a entry mass (not including the

aerobrake) of ~56.1 t. Of the total 9.9 t, ~3.9 t is

associated with the TPS system and the remaining

6.0 t with the 23 m long triconic aerobrake structure

(see Table 2). Following orbit capture, subsequent

deorbit and atmospheric reentry, the aerobrake

shell is jettisoned, and parachutes are deployed to

slow the spacecraft descent velocity to ~632 m/s.

This final terminal velocity is removed by the

descent stage which carries ~1 1 t of propellant and

uses four RL 10-class engines modified to burn

LOX/CH4. The "wet" TMI stage carries ~48 t of LH2

propellant and has a total mass of ~71.1 t resulting

in an IMLEO of ~137.1 t for the cargo lander

mission.

The ERV mission utilizes an integrated

aerobrake / hab module / TEl stage design with

LOX / CH4 engines, and has a total mass of ~74.1 t.

NASA/TM-- 1998-208834/REV 1 17
3x15 klbf ClS/NTRs

NTR TMI Stage

(Sized for 2007 ERV Cargo)

Triconic Aerobrake/Descent Shell

Containing Mars Surface Payload

(~61/74 t on Piloted/Cargo Missions)

56 t Capacity LH 2 Tank

, RCS Thrusters

4.3 m ........_ .........._,,_

7.4 m I.D.

,_, 20.6 m ......................... _i i

_ 27.5 m ....................................... _..i_ 23 m ................................

i___ 50.5 m ...................................................................................... .ii

Figure 11 .--Size, Mass and Key Features of "Common" TMI Stage and Aerobraked

Payloads.

Figure 12.--3-D Image of Expendable TMI Stage and Aerobraked Payload.

NASA/TM-- 1998- 208834/REV 1 18
Table 7. DRM "Three Mission" IMLEO Summary

("2 Perigee Burn" Earth Departure Scenario)

(IMLEO < 160 t / 2-80 t "Magnum"/Shuttle C HLLVs)

Stage/ 2011 Cargo 2011 ERV 2014 Piloted

Propulsion/Isp Element Masses (t) Lander Mission Mission Lander Mission

TEl Stage Return Habital 29.10

LOX/CH4 TEl Stage 5.89

Isp = 379 s Propellant 28.90

(O/F = 3.5:1)

Ascent Stage Crew (6) & Suits 1.44

LOX/CH4 MAV Crew Cab/ECRV 4.80

Isp = 379 s Ascent Stage 4.10

(O/F = 3.5:1) Propellant' 38.40

Descent Stage Habitat & Surface Payload 31.34 29.51

LOX/CH4 Descent Stage 4.20 4.20

Isp = 379 s Propellant** 10.98 11.38

(O/F = 3.5:1)

Aerobrake/Descent Shell 9.92 10.18 13.58

MOC System (MAB = _/Mp, (a + bV_) + Ms)+

Parachutes 0.70 0.70

Total Payload Mass 66.04 74.07 60.81

Expendable F(klbf) per eng/Isp(s) 14.76/955 14.76/955 14.76/955

TMI Stage

LH2 NTRs

@ 940-955 s

CIS Engines (#) 7.67 (3) 7.67 (3)

Radiation Shields (#)

TMI Stage

Tank & Structure 12.72 12.72 12.72

Avionics & Aux. Power 1.37 1.37 1.37

7.67 (3)

3.24 (3)

RCS Propulsion & Tankage 0.47 0.47 0.48

@ 320s

Propellants LH2 Propellant**' 47.67 52.01 48.20

/Reactants NTO/MMH Propellanl 0.77 0.77 0.88

Fuel Cell Reactants (02) 0.43 0.43 0.43

Total "Wet"+B17 TMI Stage 71.10 75.44 74.99

Total IMLEO 137.14 149.51 135.80

* Ascent propellant produced @ Mars (AV = 5625 m/s and Isp = 379 s)

** Assumes use of parachutes with descent AV = 632 m/s

*** Contains boiloff, cooldown, and "tank trapped" residuals

+ ARC Triconic aerobrake mass estimation formula (Table 2)

NASA/TM-- 1998-208834/REV 1 19
Figure 1&--Cargo Lander Showing Crew Ascent

Stage Departure.

This heavier payload increases the LH2 propellant

loading to ~52 t and the total TMI stage mass to

~75.4 t resulting in an IMLEO of ~149.5 t. The

piloted mission has an IMLEO of ~135.8 t consisting

of a 75 t TMI stage and an "integrated"

habitat / aerobrake lander configuration (shown in

Figure 14) weighing ~61 t. Approximately 31 t of

the piloted lander mass is surface payload which

includes a crew of six. Because of its fast transit

time (180 days) and higher entry velocity at Mars

(~8.7 km/s), the piloted lander also requires an

aerobrake which is ~3.5 t heavier than that used

on the preceding cargo missions. To reduce

aerobrake development costs and eliminate the

need for "customized" designs on each mission, a

"common" aerobrake configuration could be

developed and used on all cargo and piloted

Figure 14.--Piloted Lander Concept with Inflatable

Surface Habitat.

missions. The common design would be sized to

accommodate the heaviest payloads and entry

velocities anticipated over the ~15 year synodic

cycle. The use of the heavier piloted aerobrake on

the 2011 ERV mission would require enlarging the

size and propellant capacity of the TMI stage LH2

tank, further increasing the total mission IMLEO

and Magnum lift requirements.

Although a "2-perigee burn" departure scenario

has been baselined for the DRM, "single burn"

departures can also be easily accommodated on

the cargo and piloted lander missions since the

TMI stage LH2 tanks contain only ~85% of their

maximum propellant capacity. Decreasing engine

fuel temperature and Isp to 2900 K and 940 s,

and using a single burn departure increases gravity

losses, engine burn time, propellant loading and

IMLEO to ~362 m/s, 38.2 min, 52.9 t and 142.3 t,

respectively, for the cargo lander mission, and

~380 m/s, 38.7 min, 53.6 t and 141.2 t for the

piloted mission.

Following the short TMI maneuver and an

appropriate engine cooldown period, the

aerobraked payload and "spent" NTR TMI stage

separate with the Mars spacecraft continuing on its

nominal mission. The storable bipropellant RCS

system onboard the TMI stage is then used to

perform the final midcourse correction and

targeting maneuvers (AV ~100 m/s) which place

the TMI stage onto its final disposal trajectory.

Because of the miniscule burnup of enriched

uranium-235 during the Earth departure burn

(~10 grams out of 33 kilograms in each NTR core),

disposal of the TMI stage and its engines after a

single use is a costly and inefficient use of this high

performance stage. By reconfiguring the engines

for both propulsion and power generation

("bimodal" operation), a multiple burn, "power-rich"

stage with enhanced mission capabilities and

reuse potential becomes possible as we discuss

below.

"BIMODAL" NTR VEHICLE / MISSION CONCEPT

The bimodal NTR (BNTR) vehicle concept, 1°

proposed in FY93, was examined in greater detail

during this study to quantify its performance

benefits and mission versatility, and to provide a

point of comparison with the expendable TMI

stage. A "modified" DRM scenario (Figure 15) was

evaluated that employed BNTR transfer vehicles

NASA/TM-- 1998- 208834/REV 1 20
Opportunity 1 (2011)_ #"Cargo2 Return habitat, bimodal NTR

stage, backup entry capsule _MMars

"-_a__ "_'_--_'-_ /--i_Cargo I

Cargo delivered to LEO

on Magnum, rendezvous

with bimodal NTR.

Outbound Hab delivered to LEO on

(propulsive capture to

1 Sol orbit)

Ascent vehicle, prop

production, surface

exploration gear,

inflatable Hab and

surface nuclear power _.. .....

(direct entry to surface)

Opportunity 2 (2014): 1 flight " _-_

Crew of 6 aerocaptures into 1 Sol

orbit, then lands in outbound Hab.

Surface rendezvous with

Magnum. Crew of 6 delivered to LEO

in shuttle or other. Both rendezvous

with bimodal NTR.

Crew direct enters in

capsule, Apollo-style.

Earth return Hab is

discarded

Figure 15.--"Modified" Mars Mission

pre-deployed assets.

Return

(Discard one _. Crew ascends

capsule before _" return Hab

TEl) _ n capsule

J

Profile Using Bimodal NTR Vehicle Concept.

in place of the expendable TMI stage option

discussed above. A common "core" stage, used

on cargo and piloted vehicles alike, is outfitted with

three 15 klbf BNTR engines capable of providing

up to 50 kWe using any two engines. Configured

for launch on a single Magnum booster, the

bimodal core stage is not jettisoned after the TMI

maneuver but remains with the cargo and piloted

payload elements providing them with both

midcourse correction (MCC) propulsion and all

necessary power during transit. As it nears Mars,

the bimodal stage separates from the aerobraked

payload and performs its final disposal maneuvers.

A key difference between the DRM and the

bimodal option described here is the absence of the

aerobraked LOX / CH4 TEl stage which is replaced

by an "all BNTR"-powered ERV illustrated in

Figures 16 and 17. The bimodal core stage is

connected tothe hard-shelled ERV habitat module by

a rigid, spine-like"saddle truss"to which ajettisonable

"in-line" TMI propellant tank is attached. Propellant

for the Mars orbit capture MOC and TEl burns is

contained within the core stage LH 2 tank. The

554 days of contingency consumables carried by

the ERV (in case an emergency crew abort to Mars

orbit becomes necessary) is also attached to the

rear of the hab module and can be easily jettisoned

prior to TEl. In the DRM, sizeable doors must be

opened on the ERV's integrated aerobrake in order

to remove these excess consumables. Approximately 30 days after the core stage is launched, a

second Magnum booster delivers the saddle truss,

in-line propellant tank, hab module and consumables,

to LEO where rendezvous and docking with the

bimodal core stage takes place. Because of its

higher performance engines (~940 s versus 379 s

for LOX/CH 4 RL 10 engines), and the elimination

of the large 30 kWe PVA (~3.6 t) and heavy

aerobrake (~10.2 t), the BNTR / ERV is capable of a

"single burn" Earth departure while also carrying a

spare Earth return crew vehicle (ECRV) to Mars.

This enhanced vehicle capability reduces mission

risk by providing a backup option for Earth return

should a problem arise that prevents the crew

from landing on Mars and recovering their primary

ECRV from the ascent stage. Adding a spare

ECRV to the aerobraked ERV option increases its

NASA/TM-- 1998-208834/REV 1 21
"Bimodal" NTR Core

Stage w/Refrigeration

(Sized for 2016 Cargo

ERV Mission)

Refrigeration

System

3xl 5 klbf CIS/NTRs 51.0 t Capacity ,\

kWe CBC

J w/Radiator

"In-Line"

Propellant

Tank

(Tank Jettisoned)_

Strongback

Truss

29 t

LH 2 Tank '_, Capacity

'_ _LH 2 Tank

Habitat

(Payload - 18.1 t)

f

/

i

f

i

f

dettisonable i

_i_ iConsumables ! ECRV

;:, :i (-7,3 t) / (-4.8 t) /

4.3 m 2m

IMLEO: -137.7 t

Figure 16._ize, Mass and Key Features of BNTR-Powered ERV with Crew Habitat and

Spare ECRV.

Figure 17.--3-D Image of BNTR/ERV with Spare ECRV.

NASA/TM-- 1998- 208834/REV 1 22
IMLEO by an additional 10 t (from 147.5 to

~157.8 t) even using a "2 perigee burn" departure.

The bimodal core stage LH2 tank is ~19 m long

and has a maximum LH2 propellant capacity of ~51 t

using a 3% ullage factor. In addition to avionics,

storable RCS and docking systems, a turboBrayton refrigeration system is also located in the

stage forward cylindrical adaptor section to

eliminate LH2 boiloff during the lengthy (~4.2 year)

ERV mission. To remove the ~75 watts of heat

penetrating the 2 inch MLI system in LEO (where

the highest tank heat flux occurs), the Brayton

refrigeration system requires up to ~15 kWe. At the

aft end of the bimodal core stage, a conical

extension of the stage thrust structure provides

support for a "common", one-sided, pumped-loop

heat rejection radiator system. Enclosed within this

~71 m 2 conical radiator is a closed Brayton cycle

(CBC) power conversion system employing three

25 kWe Brayton rotating units (one for each

bimodal reactor) which operate at ~2/3 of rated

capacity and provide an "engine out" capability.

The turbine inlet temperature of the working gas is

~1300 K and the Brayton system specific mass is

estimated to be ~27 kg/kWe. A mass breakdown of

the common BNTR core stage used in the

"modified" DRM and the "all BNTR" mission

scenarios described below is found in Table 8.

Table 8. Mass Breakdown for "Common"

Bimodal NTR Core Sta,(

"Bimodal" NTR Core Stage Elements

Structure

Avionics and Power

Reaction Control System (RCS)

Propellant Tank (7.4 m I.D. x 19.0 m Igth.)

Passive TPS (@2" MLI)/Micrometeor Shield

LH2 Refrigeration System (@~75 Wt)

Brayton Power System (@ 50 kWe)

NTR Assemblies

• 15 klbf CIS NTRs (3)

• External Radiation Shields (3)

• Propellant Feed, etc.

Contingency (15%)

e

Mass (t)

2.5

1.47

0.45 - 0.48

5.98

1.29

0.30

1.35

6.67

0 - 2.82

0.47

3.07 - 3.50

"Dry" Bimodal Core Stage 23.55 - 26.83

LH2 Propellant (max. LH2 Capacity) 51.0

RCS Propellant 1.62 - 2.19

"Wet" Bimodal Core Stage 76.2 - 80.0

The bimodal transfer vehicle used for the cargo

lander requires a much smaller in-line propellant

tank and saddle truss arrangement (shown in

Figure 18) than that used by the "3-burn" ERV

mission, while the piloted lander requires only the

bimodal core stage (see Figure 19). Because of

the modest power needs currently identified for

the cargo lander, payload mass reductions

Figure 18.--3-D Image of BNTR Transfer Stage and Aerobraked Cargo Lander.

NASA/TM-- 1998-208834/REV 1 23
Figure 19.--3-D Image of BNTR Transfer Stage and Aerobraked Piloted Lander.

3x15 klbf CIS/NTRs

50 kWe CBC

w/Radiator

"Bimodal" NTR Core

Stage w/Refrigeration

(Sized for 2016 Cargo

ERV Mission)

Refrigeration

System

51.0 t Capacity \

LH 2 Tank \

i '\'

/ RCS _.

Jettisonable Triconic Aerobrake/Descent Shell

"In-Line" Containing Mars Surface Payload

Propellant (~56.4 - 65.0 t on

Tank Piloted/Cargo Missions)

(Optional) /

/

/

/

/ ~11 t /

Capacity Strongback i;

/

'_LH2Tank/ Truss /

'/ j

8.6 m

5.5 m

i4 27.8 m _ 7.6 m _i_ 20 m D,

'._................................................................................................................. 55.4 m ................................................................................................................. _;

IMLEO'. 1352-1500t.. .

Figure 20.--Size, Mass and Key Features of BNTR Transfer Stage for Cargo and Piloted Lander Missions.

NASA/TM-- 1998- 208834/REV 1 24
attributed to bimodal stage usage (see Figure 20)

are small (~1 t) and associated with reduced

propellant loading in the lander due to the absence

of the MCC burn. However, the bimodal stage

subsystems can support the cargo and piloted

lander missions in a number of key ways not yet

quantified. In addition to its 50 kWe power

capability, the bimodal stage's LH2 refrigeration

system can be used to eliminate boiloff from the

~4.5 t of "seed" LH2 required for ascent propellant

and water production, and its heat rejection system

can help to dissipate "decay heat" from the ~15 kWe

dynamic isotope power system (DIPS) cart used

to deploy the nuclear surface power system after

landing. For the piloted lander mission, the

elimination of the 30 kWe PVA and MCC propellant

helps to decrease descent stage propellant

requirements and aerobrake TPS mass resulting in

an ~4.5 t reduction in piloted lander mass. As with

the cargo lander mission, the bimodal stage's LH2

refrigeration system could also be used to

eliminate boiloff from the current LOX/CH4 descent

stage or a higher performance LOX / LH2 stage.

Table 9 provides an IMLEO summary for the

cargo lander, ERV and piloted lander missions

using BNTR stages, and assuming a "single burn"

Earth departure scenario. The ERV payload mass

includes a spare ECRV and 554 days of

contingency consumables (assuming 2.2 kg / day /

person and a crew of six). Because of the ERV's

lengthy mission duration and the need for multiple

engine restarts to full power, the fuel temperature

is held to 2900 K and the Isp to 940 s for

conservatism. The 79 t core stage containing ~ 50 t

of LH2 is launched on the first Magnum booster.

The second Magnum launch delivers the payload

plus the 28.4 t saddle truss and in-line tank

containing ~19.9 t of LH2. The BNTR engines

used on the cargo and piloted lander missions

are operated at the higher performance levels

(~955 s). Of the ~55.2 t of LH2 required for the

cargo mission, a minimum of ~4.2 t would be

located in the in-line tank. For the piloted lander

mission, the entire propellant load (~50.2 t) is

contained within the core stage. The total IMLEO

for this "3 mission" bimodal scenario is 422.9 t --

essentially identical to that of the DRM (Table 7)

despite the more demanding requirements levied

on the bimodal system.

A payload and stage mass comparison of the

DRM and "modified" DRM under similar operating

conditions is shown in Table 10, and Figure 21

shows the relative size and mass of the bimodal

NTR transfer vehicles used in the comparison. The

IMLEO values assume a "2-perigee burn" Earth

departure. Because the bimodal vehicles use

"standardized" components, their reduced mass

primarily reflects decreased propellant usage

during the "2 burn" TMI maneuver. For the cargo

lander mission, total propellant loading decreases

from ~55.2 t (for "single burn" departure and

Isp~955 s) to ~48.4 t (for Isp~940 s) eliminating the

need for the small in-line tank (see Figure 21). In

the case of the BNTR / ERV, the absence of the

spare ECRV further decreases propellant loading to

the point that the in-line tank is substantially offloaded--only ~42% of its maximum propellant

capacity.

Because of its higher performance and

abundant power, the BNTR / ERV mass in LEO is

~26 tons lighter than the LOX / CH4 TEl stage

which requires two large (~8 meter x 45 meter)

PVAs to provide ~30 kWe in Mars orbit. Using the

BNTR / ERV option also eliminates the development and recurring costs of the chemical TEl stage

and its 30 kWe PVA system, as well as the

recurring cost of the aerobrake needed to place

the heavy TEl stage into Mars orbit. On the cargo

and piloted hab lander missions which utilize

aerobraking, the common bimodal core stage

provides both a 50 kWe power source and the

MCC propulsion which helps reduce the size and

mass of these payload elements. Bimodal

operation also simplifies mission operations by

eliminating the need for multiple solar array

deployment / retraction cycles and the complexities

of array pointing and tracking of the Sun during

transit and while in Earth and Mars orbit. Overall, the

bimodal approach has a lower "3 mission" IMLEO

(~396 t versus 422 t for the DRM) while providing

substantially more capability. It also provides one of

the lowest cost and risk options for Mars exploration

because it requires fewer major systems.

Lastly, the requirements on total engine burn

time and fuel burnup are considered modest. For

the most demanding BNTR / ERV mission (multiple

burns and total mission duration ~4.2 years), the

total engine burn time is ~50.8 minutes, assuming

a "single burn" departure and a spare ECRV. The

TMI burn is the longest at ~36.9 minutes, and

includes the effect of a substantial gravity loss

(estimated at ~345 m/s for C3 = 8.97 km 2 / s2,

NASA/TM-- 1998-208834/REV 1 25
Table 9. Modified DRM "Three Mission" IMLEO Summary for Single Burn Earth Departure

and S )are ECRV (IMLEO < 160 t / 2 - 80 t Magnum / Shuttle C HLLVs)

2011 2011 2014

Payload/Vehicle Element Cargo Lander ERV Piloted Lander

Propulsion/Isp Masses (t) Mission Mission Mission

Earth Return Crew Hab Module 18.15

Vehicle Spare ECRV 4.83

Payload Contingency Consumables 7.31

Ascent Stage Crew (6) & Suits 1.44

LOX/CH 4 MAV Crew Cab/ECRV 4.83

Isp = 379 s Ascent Stage 4.06

(O/F = 3.5:1 ) Propellant* 38.40

Surface Payload 31.34 26.81

Descent Stage Descent Stage 4.20 4.20

LOX/CH 4 Aerobrake/Descent Shell + 9.88 13.24

Isp = 379 s Parachutes 0.70 0.70

(O/F = 3.5:1 )

Propellant** 10.03 9.99

Total Payload Mass 65.04 30.29 56.38

CIS Engines (#) 7.67(3) 7.67(3) 7.67(3)

F(klbf) per engine/Isp(s) 14.76/955 15/940 14.76/955

Radiation Shields (#) 3.24(3) 3.24(3)

"In-Line" TMI LH 2 4.26 8.52

Common Tank & Structure

NTR Vehicles TMI "Core" Stage 11.77 11.77

Tank & Structure

w/Modular TMI/MOC/TEI "Core" Stage 11.77

Components Tank & Structure

Brayton Power 1.55 1.55 1.55

CIS w/LH 2 System (@ 50 kWe)

Isp = 940 - 955 s LH 2 Refrigeration 0.34 0.34 0.34

System***

Avionics & Aux. Power 1.69 1.69 1.69

Propellant .... 55.24 69.84 50.19

RCS Propulsion & Tankage 0.54 0.56 0.54

NTO/MMH Propellant 1.89 2.19 1.83

Isp = 320 s

Total NTR Vehicle Mass 84.95 107.37 78.82

Total IMLEO 149.99 137.66 135.20

* Produced at Mars using "in-situ" resources

** Assumes parachutes and 632 m/s descent AV

*** Cooling capacity of "core" tank @ ~75 Wt

.... Contains boiloff, cooldown, "tank trapped" residual and disposal LH 2 also

+ Using ARC Triconic aerobrake mass estimation formula (Table 2)

NASA/TM-- 1998- 208834/REV 1 26
Table 10. "Three Mission" IMLEO Comparison for DRM and "Modified" DRM Using BNTR

NTR/Aerobrake (DRM) and "Modified" DRM: 80 t Magnum

Mission Feature(s): Uses JSC "Supplied" payload masses adjusted for "bimodar' NTR operation, fixed 4.2 t LOX/CH4

descent stage, 0.7 t parachutes for descent assist (,&Vdesc = 632 m/s), and "2 Perigee Burn" Earth departure.

Magnum

Launch

#1

#2

{

{

Flight Element

Mission Type

Payload

- Surface/"ln-Space"

- Transportation

"In - Line"

Propellant/Tankage

(LH 2 and/or LOX)

2011 Cargo Lander

DRM I modifiedDRM **

66.0 65.0

- 40.2 - 40.2

- 25.8 - 24.8

2011 ERV *

DRM I modifiedDRM **

74.1 25.5

- 29.1 - 25.5

- 45.0

20.8

2014 Piloted Lander

DRM I modifiedDRM **

60.8 56.4

- 30.9 - 28.4

- 29.9 - 28.0

NTR TMI stage

("Modified" DRM 71.1 73.9 75.4 79.0 75.0 75.6

uses "bimodal" NTRs)

Total: 137.1 138.9 149.5 125.3 135.8 132.0

# Magnums 2 2 2 2 2 2

Totals

modified DRM DRM **

200.9 146.9

- 100.2 - 94.1

- 100.7 - 52.8

20.8

221.5 228.5

422.4 396.2

6 6

2011 ERV mission using "bimodar' NTRs for MOC and TEl is lighter than DRM by ~24 t and eliminates DDT&E and

recurring costs for LOX/CH 4TEl stage, also recurring cost for 30 kWe PVA and aerobrake.

Common "bimodal" NTR core stage provides 50 kWe power capability to the ERV, Cargo and Piloted lander missions.

Also supplies MCC burns for these missions. For cargo lander, the "bimodar' stage refrigeration/heat rejection systems

can be used to cryocool 4.5 t of "seed" LH2 and dump "waste heat" from 15 kWe DIPS )ower cart.

_ ........................................ 28 m ......................................... _.,, ............................22.75 m ............................ _i

2011

"3 Burn" ERV

IMLEO: 125.2 t

20 m

2011 Cargo Lander

IMLEO: 138.9 t

2014 Piloted

Hab Lander

IMLEO: 132.0 t

i= 20 m

Figure 21 .--BNTR Transfer Vehicles Used in Comparison with the DRM.

NASA/TM-- 1998-208834/REV 1 27
Isp~940 s, and a vehicle thrust-to-weight ratio of

~0.15). With regard to uranium-235 consumption,

estimates indicate a fuel burnup of ~0.05% during

the "propulsion mode" and ~0.73% during the

"power mode" assuming a continuous 50 kWe

power output from the three bimodal engines over

a 5 year period.

THE "ALL PROPULSIVE" BIMODAL NTR OPTION

The next logical application of the BNTR stage

beyond the modified DRM is propulsive capture of

all payload elements into Mars orbit. This "all

propulsive" NTR option makes the most efficient

use of the bimodal engines which are now available

to supply abundant power to spacecraft and

payloads in Mars orbit for long periods. Even after

payload separation and landing on the Mars

surface, the core stages become valuable orbiting

resources and can serve as high power communications relays and/or surface navigational aids.

Propulsive capture into the reference "250 km by

1 sol" ellpitical Mars parking orbit also makes it

possible to design a standardized, reduced mass

"aerodescent" shell for landing all payloads on the

Mars surface. From this reference parking orbit

(similar to that used by the Viking lander missions in

1976), the payload entry velocity is ~4.5 km/s and

the mass of the "triconic-shaped" aerodescent

shell varies by only ~0.53 t over a payload mass

range of 40 to 65 t (using equation in Table 2).

The size, mass and key vehicle features for the

"all BNTR" Mars mission option is shown in

Figure 22 and the associated cargo lander, ERV and

piloted lander IMLEO values are summarized in

Table 11. With propulsive capture, the total cargo

lander mass decreases from ~66 t in the DRM to

~62.3 t, which is attributed to a lighter aerodescent

shell (~8.2 t) and a reduced descent stage

propellant loading (~8.9 t). A detailed "3 mission"

IMLEO summary for the "all BNTR" option is found

in Appendix Table A-4. Despite this mass reduction, the substantial payload carried by the cargo

lander increases the propellant requirements on

the BNTR transfer vehicle to ~68.3 t with the core

stage holding 51 t. The remaining ~17.3 t of LH2 is

contained in the common 11.5 m in-line tank also

used on the ERV and piloted lander missions. The

total mass of the "in-line" tank, its propellant load

3x15 klbf CIS/NTRs

/

i

i 50 kWe CBC

! w/Radiator

"Bimodal" NTR Core

Stage w/Refrigeration

(Sized for 2016 Cargo

ERV Mission)

Refrigeration

System

51.0 t Capacity

LH2 Tank ,,

"In-Line"

Propellant

Tank

(Tank Jettisoned)

Strongback

29.0 t Truss

Capacity /

/

g

i

Triconic Aerodescent Shell

Containing Mars Surface Payload

(~48.7 - 62.3 t for

Piloted/Cargo Missions)

/

/

/

......... 3.5 m

ii

ii

8.6 m

20 m

63 m

-\

IMLEO" 163.1 - 164.3 t

Figure 22.--Characteristics of "All BNTR" Transfer Vehicles for Piloted and Cargo Lander Missions.

NASA/TM-- 1998- 208834/REV 1 28
Magnum

Launch

Table 11. Payload/Stage Mass Manifest for "All BNTR" Option

2011 Cargo

Lander*

2011

ERV

2014 Piloted

Flight Element Lander* Totals

Payload 62.3 25.5 48.7 136.5

- Surface/"ln-Space" - 40.2 - 25.5 - 28.0 - 93.7

- Transportation - 22.1 - 20.7 - 42.8

"In - Line"

Propellant/Tankage 25.8 20.8 35.0 81.6

(LH2 and/or LOX)

"Bimodar' NTR

Core Stage 76.2 79.0 79.4 234.6

Total : 164.3 125.3 163.1 452.7

# Magnums 2 2 2 6

Common "bimodal" NTR core stage provides 50 kWe power capability to the ERV, Cargo and

Piloted lander missions. Also supplies MCC burns for these missions. For cargo lander, the

"bimodar' stage refrigeration/heat rejection systems can be used to cryocool 4.5 t of "seed" LH2

and dump "waste heat" from 15 kWe DIPS power cart.

and the cargo lander determine the maximum lift

requirement for the Magnum booster which is ~88 t

for this mission option. Because the maximum

possible payload length for the Magnum booster is

~33 m (including the 28 m cylindrical section and

payload shroud nose cone), a smaller in-line tank

or shortened triconic aeroshell length (to ~18 m) is

required to launch these components on a single

88 t Magnum.

The piloted lander mission employs a "220 day"

outbound transit time (C3 = 14.47 km 2 / s2) to Mars

to maintain LH2 propellant requirements within the

maximum propellant capacity provided by the

common vehicle design. A "2-perigee" burn Earth

departure is also assumed for all three missions to

reduce gravity losses. With propulsive capture, the

total piloted lander mass is decreased by ~20%

(from ~61 t down to ~49 t). The main reductions are

in the aerodescent shell mass (~7.9 t versus 13.6 t

for the DRM) and the reduced descent stage

propellant loading (~7.9 t compared to 11.4 t in the

DRM). The piloted mission has longest total

engine burn time at ~58 minutes. This includes 45

minutes for the 2 perigee burns, ~12 minutes for

MOC, and an ~1 minute disposal burn to remove

the bimodal core stage from Mars orbit after crew

departure and send it into heliocentric space

(see Table 5).

the lightweight, inflatable "TransHab" module. 12

TransHab was designed to be launched in the

Space Shuttle cargo bay fully outfitted. A central

structural core ~3.4 m in diameter provides

regenerative life support, thermal control, crew

accommodations, avionics and communications,

meteoroid and orbital debris protection, a storm

shelter for crew radiation protection, and an airlock.

Once on orbit, the outer shell surrounding the

central core is inflated and corrugated flooring and

partitions are deployed into place. Fully inflated,

TransHab has an outer diameter of ~9.44 m, a

height of ~9.65 m, and provides ~500 m 3 of

habitable volume (see Figure 23).

"ALL BNTR" OPTION USING "TRANSHAB"

The attractiveness of the "all propulsive" bimodal

NTR option is further increased by the utilization of

Figure 23.--Illustration Showing TransHab Module

Attached to International Space Station.

NASA/TM-- 1998-208834/REV 1 29
"Bimodal" NTR Core "In-Line"

Stage w/Refrigeration Propellant

(Sized for 2016 Cargo Tank

ERV Mission) (Tank Jettisoned)

i

Refrigeration ) Strongbackl

System i Truss

3x15 klbf CIS/NTRs 51.0 t Capacity ,, 29 t

! LH 2 Tank _ Capacity /

/ 50 kWe CBC _' _,i ILH 2 Tank /

i w/Radiator _i ,_ _ ,

i ,,: i RCS..,. _ ) i

"TransHab" Module

(Payload - 15.0 t)

Jettisonables

Consumables

(~7.1 t)

!

/

9.44 m

-il-2m

1-_ 9.65m _-

i_,.......................................... 28 m .......................................... m-)_.................................. 256 m .................................... _,

i" ........................................................................................... 53.6 m .......................................................................................... _-

I IMLEO: ~139.3

Figure 24.--Size, Mass and Key Features of Reusable Bimodal ERV Using TransHab.

Figure 25.--3-D Image of Reusable Bimodal ERV with Inflatable TransHab Crew Module.

NASA/TM--1998-208834/REV1 30
In addition to volume augmentation, the

substitution of TransHab for the heavier, hard-shell

hab module used on the bimodal ERV in

Figures 16 and 17, provides an ~18% reduction in

element mass and introduces the potential for propulsive recovery of the bimodal ERV in Earth orbit

and its reuse on subsequent missions. The characteristics and 3-D image of the reusable bimodal ERV

and TransHab crew module are shown in Figures 24

and 25, respectively. The reusable ERV departs

Mars on February 7, 2016 and returns to Earth

180 days later on August 5, 2016. The crew

reenters directly using the ECRV, while the ERV

propulsively captures into a 500 km by ~71,136 km

elliptical parking orbit with a period of ~24 hours.

Using a 2-perigee burn departure, the reusable

ERV mission utilizes nearly the full propellant

capacity of the bimodal core stage and its in-line

tank. At a hydrogen exhaust temperature of

~2900 K (Isp~940 s), the bimodal engines are

estimated to have a "full power" operational lifetime

of ~4.5 hours. With a total engine burn time of

~58 minutes for the four primary maneuvers, a multimission capability exists for the bimodal ERV. At

Earth, a space-based upper stage could rendezvous with the ERV supplying it with a small in-line

tank containing the propellant needed to return

the ERV to LEO. Here, the core stage could be

refueled, a new in-line propellant tank attached,

and necessary consumables provided for the next

mission. Reuse of the core stage, saddle truss

and TransHab would reduce vehicle recurring

costs but must be evaluated against the increased

development and operational costs of the support

infrastructure.

Although the diameter of the aerodescent shell

does not allow the same degree of volume augmentation available on the ERV mission, the use of

TransHab on the piloted lander reduces its mass

and allows the inflatable surface hab module on the

cargo lander to be offloaded to the piloted mission.

This and a 210 day outbound transit time results

in a total propellant requirement of ~79.2 t with

~28.2 t located in the in-line tank. It is the

combined "wet" in-line tank and piloted lander

mass that sizes the Magnum lift capability at ~85 t.

By offloading the inflatable surface hab from the

cargo lander, the propellant loading in the bimodal

transfer vehicle is also reduced to ~65.5 t. The

payload and stage mass manifest for the two cargo

and piloted flights are summarized in Table 12. The

IMLEO values for the two lander missions reflect a

2-perigee burn departure and engine operation at

an Isp value of 955 s.

Table 12. Payload/Stage Mass Manifest for "All BNTR" Option Using TransHab

Mission Feature(s): "Bimodar' NTR Core Stage provides power, MCC and all primary propulsion. ERV

propulsively returned to Earth orbit. JSC "TransHab" masses for piloted lander and ERV. Fixed 4.2 t LOX/CH4

descent stage and 0.7 t parachutes for descent assist (AVdesc = 632 m/s). Inflatable surface hab module (-3.1 t) is

"offloaded" from the cargo to the piloted lander mission.

Magnum

Launch

#1

2011 Cargo

Lander*

2011

ERV*

2014 Piloted

Flight Element Lander* Totals

Payload 58.5 22.0 47.9 128.4

- Surface/"ln-Space" - 37.1 - 22.0 - 27.3 - 86.4

- Transportation - 21.4 - 20.6 - 42.0

"In - Line"

Propellant/Tankage 23.0 37.3 36.7 97.0

(LH2 and/or LOX)

"Bimodar' NTR

Core Stage 76.1 80.0 79.3 235.4

Total : 157.6 139.3 163.9 460.8

# Magnums 2 2 2 6

Common "bimodar' NTR core stage provides 50 kWe power capability to the ERV, Cargo and

Piloted lander missions. Also supplies MCC burns for these missions. For cargo lander, the

"bimodar' stage refrigeration/heat rejection systems can be used to cryocool 4.5 t of "seed" LH2

and dump "waste heat" from 15 kWe DIPS power cart.

NASA/TM-- 1998-208834/REV 1 31
MARS/PHOBOS MISSION OPTION USING LANTR

The benefits of a human expedition to Phobos

have been discussed previously 2,25 and range

from basic scientific knowledge to practical applications of the moon as an operating node and

potential propellant depot for future human

exploration and development activities on Mars.

The Mariner 9 and Viking Orbiter missions in the

1970s provided images and spectral data

suggesting that both Phobos and Deimos were

formed within the asteroid belt and later captured

by Mars. Their low mean densities (~2 g/cm 3) and

reflectivities 26also suggest a chemical composition

similar to carbonaceous chondrite meteorites,

which contain substantial quantities of water and

carbon-containing materials. Should this be true,

Phobos could provide an abundant source of

propellants for future reusable Mars transfer and

landing vehicles. A Phobos mission would also

provide expertise on operations both near and on

a small, essentially gravity free planetary body of

value to the exploration of other near Earth

asteroids.

The introduction of LANTR and its integration

into the bimodal stage opens the possibility for a

"side trip" to Phobos within the current DRM. The

reusable ERV mission just discussed showed the

benefits of using TransHab. It also indicated, however, that the second Magnum booster was only

utilizing ~75% of its lift capability in launching

TransHab, the in-line propellant tank and saddle

truss (see Table 12). Stretching the in-line LH2 tank

size and propellant capacity is also limited because

of the volume constraints of the Magnum payload

shroud. Using LANTR engines at modest O/H

mixture ratios increases bulk propellant density (by

substituting high-density LOX for low-density LH2)

and improves vehicle performance while staying

within the available payload length limits. LANTR

operation also helps to increase engine thrust,

shorten burn times and extend engine life.

Phobos Mission Description Usinq LANTR

The Phobos mission scenario utilizes LANTR

engines only for Earth departure. At an operating

temperature of 2900 K and an O/H MR = 0.0 (LH2

only operation), the thrust from the LANTR engine

is 15 klbf (see Figure 8). At an O/H MR = 0.5, the

thrust per engine is increased by a factor of ~1.33

while the Isp decreases from ~940 s to 831 s.

During the ~29 minute long, 2-perigee burn TMI

maneuver, the three LANTR engines produce a

total thrust of ~59.7 klbf while using ~39.5 t of LH2

(including "cooldown" propellant) and ~19.2 t of

LOX. Following the TMI burn, the spent in-line LH2

tank and two spherical LOX tanks attached to it are

jettisoned from the saddle truss to reduce vehicle

weight. On all subsequent burns, the LANTR

engines operate at MR = 0.0 and Isp = 940 s. The

bimodal LANTR vehicle concept with TransHab

crew module is illustrated in Figure 26 and its

corresponding 3-D image is shown in both

Figure 27 and on our cover page.

At Mars, the LANTR transfer vehicle propulsively

captures into a 250 km by 33,793 km elliptical

parking orbit where it remains during most of the

crew surface stay. Approximately 32 days before

TEl, the LANTR ERV jettisons its ~6.3 t of

contingency consumables and then executes

three propulsive maneuvers to rendezvous with

Phobos. At apoapse, the LANTR engines burn to

change plane to near equatorial. The required AV

is ~212 m/s assuming an arrival declination of

~27 degrees. Next, the periapse is raised to Phobos

altitude of 5981 km (AV ~228 m/s). A final

circularization burn to lower apoapsis to 5981 km

requires a AV of ~664 m/s. Including an additional

~100 m/s to rendezvous with Phobos, the total AV

requirement is ~1105 m/s.

Once in position, the crew lifts off from the Mars

surface and rendezvouses with the LANTR / ERV

to begin a month long investigation of Phobos.

Detailed spectroscopic analysis and other scientific

measurements (including impact probes and deep

penetrating radar imaging) would be carried out

onboard the ERV to determine whether or not

water is present. Prior to TEl, the ERV departs its

near-equatorial orbit and returns to an inclined

elliptical orbit matching the declination for the

outgoing launch asymptote. The same ~1105 m/s

is assumed for these return maneuvers. The total

IMLEO for the LANTR / ERV mission to Phobos is

~157.9 t with each Magnum booster now delivering ~79 t to LEO (see Table 13). The cargo lander

mission is unchanged from Table 12 and the

piloted lander mass decreases slightly due to the

shortened surface stay time (~475 days) and

reduced crew consumables required for the

Phobos mission.

By stretching the LANTR / ERV in-line LH2 tank

size and capacity to ~13.5 m and 35 t to increase

performance, a more robust Phobos exploration

scenario is possible. Rather than relying on remote

data acquisition alone, the "stretch" LANTR / ERV

NASA/TM-- 1998- 208834/REV 1 32
"Bimodal" NTR Core

Stage w/Refrigeration

(Sized for 2016 Cargo

ERV Mission)

Refrigeration

3x15 klbf CIS/LANTRs System

! 51.0 t Capacity

i LH 2 Tank

150 kWe CBC ! _i.

w/Radiator

"In-Line"

Propellant

Tank

(Tank Jettisoned)

Strongback

Truss

-10 t ! Jettisonable

Capacity Consumables ! 29 t

LOX Tanks _i Capacity (~6.3 t) /

/(2) ! LH 2 Tank

"TransHab" Module

(Payload - 15.4 t)

!

2m

IMLEO: ~157.9t }

Figure 26.--Size, Mass and Key Features of Bimodal LANTR Transfer Vehicle for Mars/Phobos

Mission Option.

Figure 27.--3-D Image of Bimodal LANTR Transfer Vehicle for Mars/Phobos Mission Option.

NASA/TM-- 1998-208834/REV 1 33
Table 13. Payload/Stage Mass Manifest for Bimodal "LANTR" Mars/Phobos Option

Mission Feature(s): Bimodal "LANTR"-powered ERV visits Phobos before TEl. LANTR engines provide thrust

augmentation (MR = 0.5) for TMI with MR = 0 for remaining primary propulsive maneuvers. "TransHab" mass

used on ERV and piloted mission. Fixed 4.2 t LOX/CH, descent stage and 0.7 t parachutes for descent assist

(AVd_,o= 632 m/s).

Magnum

Launch

#1

2011 Cargo

Lander*

2011 ERV*

(Visits Phobos)

2014 Piloted

Flight Element Lander* Totals

Payload 58.5 21.7 47.0 127.2

- Surface/"ln-Space" - 37.1 - 21.7 - 26.6 - 85.4

- Transportation - 21.4 - 20.4 - 41.8

"In - Line"

Propellant/Tankage 23.0 57.0 35.8 115.8

(LH2 and LOX)

"Bimodar' NTR

Core Stage 76.1 79.2 79.3 234.6

Total : 157.6 157.9 162.1 477.6

# Magnums 2 2 2 6

* Common "bimodar' NTR core stage provides 50 kWe power capability to the ERV, Cargo and

Piloted lander missions. Also supplies MCC burns for these missions.

(shown in Figure 28) would carry a 2-person

"multiple sortie" lander and ~250 kg of scientific

equipment to Phobos orbit. The ~6.3 t of

contingency consumables are also transported to

Phobos orbit to build up an easily accessible

emergency food cache thereby allowing

subsequent missions to transport an inflatable

surface hab and other equipment needed to

establish a permanent foothold on Phobos. The

Phobos lander (shown to scale in Figure 28) is

sized for ten round trip sorties to the surface of

Phobos and back. On each mission, two

astronauts deploy ~25 kg of scientific equipment

and return to the ERV with ~10 kg of samples.

Because the escape velocity from Phobos is very

low (~15 m/s), the total storable propellant

requirements for the entire ten mission set is only

~160 kg. The ~1.73 t Phobos lander mass

includes the "dry" lander (at ~1.10 t) and its

propellant load (~0.16 t), two EVA suits with life

support (~0.22 t) and scientific equipment

(~0.25 t). The payload / stage mass manifest for

this robust Phobos option is provided in Table 14

and the associated "3 mission" IMLEO summary in

Table 15. To compensate for the increased propellant Ioadings in the in-line LH2 and LOX tank sets,

the TransHab crew module and 32 days of extra

consumables (totaling ~15.4 t) are delivered to the

ERV using the Space Shuttle or "lower cost" RLV.

The remaining ~155.6 t are launched on two

Magnums.

AN "ALTERNATIVE MISSION PROFILE" USING

BNTR AND TRANSHAB

The BNTR transfer vehicle in combination with

TransHab provides a high degree of mission

versatility. In addition to providing a reuse capability

for the ERV, a Phobos mission option is also

possible through the addition of LOX "afterburner"

nozzles and propellant feed system for LANTR

operation. The BNTR and TransHab combination

also allows one to consider an alternative mission

profile in which the crew travels to and from Mars

on the same bimodal transfer vehicle as depicted

schematically in Figure 29. This approach cuts the

duration of the ERV mission approximately in half--

from ~4.7 to 2.5 years--while the remaining two

mission elements (the cargo and "unpiloted" crew

lander) are left unattended by humans for no more

than ~2.8 years.

The roundtrip piloted transfer vehicle departs

Earth on January 21,2014 (C3 = 15.35 km 2 / S2)

and propulsively captures into Mars orbit 210 days

later on August 19, 2014. The outbound transit

time is extended by 30 days to maintain propellant

requirements within the capacity of the bimodal

NASA/TM-- 1998- 208834/REV 1 34
3x15 klbf ClS/LANTRs

i

I

/

"Bimodal" NTR Core

Stage w/Refrigeration

(Sized for 2016 Cargo

ERV Mission)

RefrigerationS,

System

"In-Line"

Propellant

Tank

(Tank Jettisoned)

Strongback

Truss

-11 t /

"TransHab" Module

(Payload - 15.4 t)

Jettisonable

51.0 t Capacity :', Capacity i 34.8 t Consumables

LH 2 Tank ', (~6.3 t)

50 kWe CBC _ LOX Tanks _ Capacity

w/Radiator i (2) ! LH 2 Tank /'

*t

Phobos

Lander

(-1.7 t)

i

9.44 m

4.5 m

3.5 m

i-_......................... 19 m

28 m

•..........2.7 m I.D. 2m

59.4 m

_" MLEO: 'i71t _i

Figure 28.--Size, Mass and Key Features of "Stretch" LANTR/ERV for Phobos Lander Option.

Table 14. Mass Manifest for "Stretch" LANTR/Phobos Lander Mission

Mission Feature(s): Bimodal "LANTR"-powered ERV visits Phobos before TEl. LANTR engines provide thrust

augmentation (MR = 0.5) for TMI with MR = 0 for remaining primary propulsive maneuvers. "TransHab" mass used on

ERV and piloted mission. Fixed 4.2 t LOX/CH4 descent stage and 0.7 t parachutes for descent assist (AVdesc = 632 m/s).

ERV also carries "2 person" multiple sortie Phobos lander and scientific equipment.

Magnum

Launch

#1

#2

2011 Cargo

Lander*

2011 ERV*

(Visits Phobos)

2014 Piloted

Flight Element Lander* Totals

Payload 58.5 23.4 47.0 128.9

- Surface/"ln-Space" - 37.1 - 23.4 - 26.6 - 87.1

- Transportation - 21.4 - 20.4 - 41.8

"In - Line"

Propellant/Tankage 23.0 68.4 35.8 127.2

(LH2 and LOX)

76.1 79.2 79.3 234.6

Total ' 157.6 171.0 ÷ 162.1 490.7

# Magnums 2 2 + 2 6

"Bimodal" Core StageNTR

* Common "bimodal" NTR core stage provides 50 kWe power capability to the ERV, Cargo and Piloted

lander missions. Also supplies MCC burns for these missions. For cargo lander, the "bimodar' stage

refrigeration/heat rejection systems can be used to cryocool 4.5 t of "seed" LH2 and dump "waste heat"

from 15 kWe DIPS power cart.

+ On 2011 ERV mission, "TransHab" module and extra consumables (-15.4 t) would be launched on

Shuttle or lower cost RLV with remaining mass (-155.6 t) launched on two Magnums.

NASA/TM-- 1998-208834/REV 1 35
Table 1 5.

Payload/Vehicle

Propulsion/Isp

Earth Return

Vehicle

Payload

Ascent Stage

LOX/CH4

Isp = 379 s

(O/F = 3.5:1 )

Descent Stage

LOX/CH 4

Isp = 379 s

(O/F = 3.5:1 )

Common

NTR/LANTR Vehicles

w/Modular

Components

LH2 NTR

Isp = 940 s

LANTR

Isp = 831s @ MR=0.5

for TM I

RCS

NTO/MMH

Isp = 320 s

IMLEO Summary for Phobos Lander Option Using

("Single Burn" Earth De 3arture Scenario)

(IMLEO < 166 t/2-83 t Ma,(.num/Shuttle C HLLVs)

Element

Masses (t)

'q-ransHab" Modulet

Extra Consumablest

Contingency Consumables

Phobos Lander

& Science Equipment

Crew (6) & Suits

MAV Crew Cab/ECRV

Ascent Stage

Propellant*

Surface Payload

Descent Stage

Aerodescent Shell +

Parachutes

Propellant**

Total Payload Mass

NTR/LANTR Engines (#)

2011

Cargo Lander

Mission

F(klbf) per engine/Isp (s)

Radiation Shields (#)

"In-Line" TMI LH 2

Tanks & Structure

"In-Line" TMI LOX

Tanks & Structure

TMI "Core" Stage

Tanks & Structure

TMI/MOC/TEI "Core" Stage

Tanks & Structure

Brayton Power

System (@ 50 kWe)

LH 2 Refrigeration System***

Avionics & Aux. Power

LH 2 Propellant ....

LOX Propellant

Propulsion & Tankage

Propellant

2014

Piloted Lander

Mission

Total NTR Stage Mass

Total IMLEO

1.44

4.83

4.06

38.40

28.24 25.14

4.20 4.20

8.15 7.90

0.70 0.70

8.30 7.62

58.48 47.00

7.67(3) 7.67(3)

14.76/955 14.76/955

3.24(3)

8.25 8.25

11.77 11.77

1.55 1.55

0.60 0.60

1.69 1.69

65.54 78.26

0.52 0.52

1.55 1.54

99.14 115.09

157.61 162.09

2011

ERV

Mission

14.96

0.42

6.27

1.73

23.38

8.13(3)

19.9/831

15/940

3.24(3)

9.88

0.49

11.77

1.55

0.34

1.69

85.34

22.20

0.57

2.38

147.58

170.96

t Delivered on Shuttle or lower cost RLV

* Produced at Mars using "in-situ" resources

** Assumes parachutes and 632 m/s descent AV

*** Cooling capacity of "core"/"in-line" tank @ -75/46 W t, respectively

.... Contains boiloff, cooldown, "tank trapped" residual and disposal LH2 also

+ Using ARC Triconic aerobrake mass estimation formula with Vo=4.5 km/s

NASA/TM-- 1998- 208834/REV 1 36
Cargo/hab lander and "in-line"

tanks delivered to LEO on

to LEO in shuttle or other. Both

rendezvous with bimodal NTR.

Opportunity I (2011): 2 flights _ Cargo 2

Lander/Habitat, consumables, rover

(propulsive capture to I Sol orbit) __-" ill-'Mars

Ascent vehicle, Pr°P_ ;_@'_-'t_" A

production, surface [ -- _ Cargo

exploration gear, | / I \

inflatable Hab and

Magnum, rendezvous surface nuclear power ,jDII_... ........

with bimodal NTR. (propulsive capture to ,/ "_

1 Sol orbit) (/_

Opportunity2 (2014):1 flight ___// /

..... Crew of 6 propulsively captures into 1 Sol _ IJ I l}1_

................. _........ orbit, rendezvous with orbiting hab lander, _ _ _

iranst_aD/ in-nine tanK aeiive..red to descend to surface. TransHab remains in _ / ......

LEO on Magnum. Crew of 6 delivered orbit for subsequent crew return.

Crew direct enters in :

capsule, Apollo-style. :

TransHab is discarded

,,'X_" "C;'rewascends

(Retain crew capsule J /_ to TransHab

for reentry) ..__/ in capsule

I

Figure 29.--"Alternative Mission Profile": Round Trip Piloted Transfer Vehicle Using BNTR and TransHab.

core stage (-51 t) and its 11.5 m "in-line" tank

(-29 t). Table 16 shows the outbound piloted

transit times possible over a 15 year period using

the common bimodal transfer vehicle. Return

transit times are held constant at 180 days.

Once in Mars orbit, the crew transfer vehicle

(CTV) rendezvouses with the "unpiloted" hab

lander (which is now delivered on an earlier cargo

mission) and then descends to the surface. The

absence of crew on the hab lander mission

eliminates the need for 210 days of outbound consumables (-2.77 t) and the engine crew radiation

shields (-3.24 t). This allows the hab lander to carry

the inflatable surface module (-3.1 t) and science

equipment (-4.4 t) previously carried on the

crowded cargo lander. Size, mass and key features

of the bimodal vehicles used on the piloted

and cargo / hab lander missions are shown in

Figures 30 and 31, respectively.

The piloted transfer vehicle uses the same

common core stage, in-line propellant tank and

saddle truss utilized on the bimodal E RVs discussed

previously. The TransHab payload mass

(-16.8 t) includes the mass of the six crew and their

suits, and 30 days of extra consumables to account

for the longer outbound transit time. Contingency

consumables (-6.7 t) consistent with a 507 day

surface stay are also carried. The total propellant

required for the mission is -79 t, and the total

vehicle length and IMLEO are -54 m and -140 t,

respectively. A smaller (-6.5 m), in-line propellant

tank is used on the common bimodal transfer

vehicles that deliver the -46 t hab and -54 t cargo

landers into Mars orbit. The total propellant needs

for these transfer vehicles are -57.3 t and -64.3 t,

respectively. A 3-D image of the bimodal cargo

transfer vehicle showing its relative size is shown in

Figure 32, and Table 17 summarizes the payload /

stage mass manifest for the "3 mission" set.

A detailed IMLEO summary is found in Appendix

Table A-5.

SUMMARY COMMENTS AND DISCUSSION

The bimodal NTR propulsion and power system

provides an extremely versatile space transportation option to the planners and designers of

future human exploration missions to Mars.

NASA/TM-- 1998-208834/REV 1 37
¢-

O

"iD

o

121 ._

"iD e,

Q

E

I-"

Q.

I---

¢-

O

<1

"0

,,_

0

O

c6

.Q

I-"

NASA/TM--1998-208834/REV1 38
3x 15 klbf CIS/NTRs

i

/ 50 kWe CBC

/ w/Radiator

4.3 m

"Bimodal" NTR Core "In-Line"

Stage w/Refrigeration Propellant

(Sized for 2016 Cargo Tank

ERV Mission) (Tank Jettisoned)

Refrigeration

System Strongback 51.0 t Capacity 29 t Truss

LH2 Tank ', Capacity /

/ _, LH 2 Tank

RCS-.. _ " / /

"TransHab" Module

(Payload 16.8t)

Jettisonable

Consumables

(~6.7 t)

//

/

9.44 m

19m

28 m

IMLEO: -139.6 t

Figure 30.--Size, Mass and Key Features of Round Trip Piloted Transfer Vehicle.

3x15 klbf CIS/NTRs

"Bimodal" NTR Core "In-Line"

Stage w/Refrigeration Propellant

(Sized for 2016 Cargo Tank

ERV Mission) (Tank Jettisoned)

Refrigeration

System Strongback!

51.0 t Capacity Truss

LH2 Tank 14.0t i

! 50 kWe CBC i !_ Capacity '_

i w/Radiator i :.LH2 Tank !i

i8.6m

i E_ 3 5 m

4.3m i_. ..! " , .+

27.8m - gm ,_ 20m -

i" 56.8 m "_i

IMLEO: ~132.9 - 148.2 t

Figure 31 .--Size, Mass and Key Features of BNTR Vehicles for Cargo and "Unpiloted" Hab Lander

Missions.

Triconic Aerodescent Shell

Containing Mars Surface Payload

(~45.8 - 53.7 t for Unpiloted/Cargo

Lander Missions)

NASA/TM-- 1998-208834/REV 1 39
Figure 32.--3-D Image Showing Relative Size of Bimodal Cargo Transfer Vehicle.

Table 17. Payload/Stage Mass Manifest for Alternative Mission Option

Mission Feature(s): Crew travels to and from Mars using "bimodal" NTR transfer stage and "TransHab."

Results based on JSC "supplied" payload masses adjusted for "bimodal" NTR operation, fixed 4.1 t

LOX/CH4 descent stage and 0.7 t parachutes for descent assist (AVdesc = 632 m/s). A "Single Burn" Earth

departure is used along with outbound/inbound transit time of 210/180 days, respectively.

Magnum

Launch

2011 Cargo

Lander**

2011

"Unpiloted"

Hab Lander**

2014 Piloted

Flight Element Mission* Totals

Payload 53.7 45.8 23.5 123.0

- Surface/"ln-Space" - 33.3 - 25.4 - 23.5 - 82.2

- Transportation - 20.4 - 20.4 - 40.8

"In - Line" Propellant

Tankage/Structure 18.4 11.2 37.1 66.7

(LH2 and/or LOX)

"Bimodal" NTR

Core Stage 76.0 75.8 79.0 230.8

Total : 148.1 132.8 139.6 420.5

# Magnums 2 2 2 6

* 2014 Piloted "round trip" transfer vehicle uses "bimodar' NTRs for MOC and TEl also, and

eliminates the DDT&E and recurring costs for the LOX/CH4 TEl stage, as well as recurring

cost for the 30 kWe PVA and aerobrake.

** Common "bimodar' NTR transfer stage also provides 50 kWe power capability to the cargo

and "unpiloted" Hab lander missions. Also supplies MCC burns for these missions.

NASA/TM-- 1998- 208834/REV 1 40
Bimodal operation fully exploits the true

performance potential of the NTR by tapping into

the "rich source of energy" that exists within the

engine's reactor core. Rather than throwing away a

valuable transportation system asset after a single

use, "better systems engineering" has led to the

design of an integrated NTR "core" stage providing

both propulsion and power generation. The core

stage uses three small (~15 klbf) bimodal NTR

engines providing up to 50 kWe of electrical

power, a portion of which (~15 kWe) is used to

support an active refrigeration system for "zeroboiloff," long term storage of LH2 propellant. The

bimodal stage uses a Brayton power conversion

system enclosed within the vehicle's conical thrust

structure which also provides support for a

common heat rejection radiator system. The

incorporation of power generation and refrigeration

systems results in a smaller, higher performance

NTR stage with multiple burn, propulsive capture

and reuse features. The use of multiple small

engines also provides an "engine out" capability

for the vehicle and should aid in the design of

"contained" ground facilities for rigorous engine

testing that are both cost-effective and meet

current environmental regulations.

A simpler, lower cost transportation system

requiring fewer major elements and providing

greater mission capability are a few of the major

benefits of the bimodal NTR option. Table 18

compares and summarizes the number of mission

elements and the ETO requirements for the DRM,

"modified" DRM and "all BNTR" options examined

in this paper. The DRM uses NTR propulsion for

TMI, a large 30 kWe PVA for in space power,

a heavy, common aerobrake/descent shell for

MOC and reentry, a "SP-100" type nuclear reactor

for surface power, LOX/CH4 engines for TEl and an

ECRV for Earth return--a total of 6 mission

elements. The introduction of the BNTR in the

"modified" DRM cuts this number in half

(lowering DDT&E and recurring costs) while

increasing the available power to payloads in

transit and in Mars orbit to 50 kWe. The use of

standardized modular components in the bimodal

Table 18. Corn

Mission Elements

and ETO Requirements

TMI

In-Space Power

MOCS

Mars Orbit Power

Mars Reentry

System

Surface Power

TEl

EOC

Total # Major Systems

# Magnum Launches

[Required lift (t)]

IMLEO (t)

)arison of DRM, "Modified" DRM, and "All BNTR" Mars Mission Options

DRM

NTR

PVA

(30 kWe)

AB +

PVA

(30 kWe)

Common

AB/AS ++

Nuc. Rx.

(Brayton)

LOX/CH4

ECRV t

"Modified"

DRM

(BNTR)

BNTR *

BNTR

(50 kWe)

AB & BNTR

BNTR

(50 kWe)

Common

AB/AS

Common Rx

(Brayton)

BNTR

ECRV

3

6

[80]

- 396

6

6

[8O]

- 422

"All NTR"

(BNTR)

BNTR

BNTR

(50 kWe)

BNTR

BNTR

(3x50kWe)

AS

Common Rx.

(Bra_on)

BNTR

ECRV

3

6

[88]

-453

"All NTR"

(BNTR) with

"TransHab"

BNTR

BNTR

(50 kWe)

BNTR

BNTR

(3x50kWe

AS

Common Rx

(Bra_on)

BNTR

ECRV &

BNTR

3

6

[85]

- 461

"All NTR"

(BNTR) with

"TransHab"

and LANTR

BLANTR**

BLANTR

(50 kWe)

BLANTR

BLANTR

(3x50kWe)

AS

Common Rx.

(Bra_on)

BLANTR

ECRV

4

6

[83]

- 478-491

ALT. ARCH.

"All BNTR"

with

"TransHab"

BNTR

BNTR

(50 kWe)

BNTR

BNTR

(3x50kWe)

AS

Common Rx.

(Bra_on)

BNTR

ECRV

3

6

[8O]

- 421

* BNTR: "Bimodar' NTR with Brayton Power Conversion/** BLANTR: BNTR with "LOX Afterburner" Nozzle

÷Aerobrake/÷÷Aerodescent shell/1- ECRV: Earth Crew Return Vehicle

NASA/TM-- 1998-208834/REV 1 41
2011

Earth Return Vehicle

.......... .............. iii 

..... _iiii!!!!!!!!!i!!!!!::!::_::_::_::_::_::_::_::_::_::i::i::i::iii::i::i::iii:_ii?):` ::_NNii_i iiii_ii?:i::i::i::i!iii ii?iiiiiiiiI :ri :_l

i"_iiiiiiiiiiiiiiiiiiiiiiiiii_

2011 .iiii.!__ ................

Cargo Lander

2014

Piloted Lander

Figure 33.--Family of Bimodal NTR Transfer Vehicles Using "Modular" Components.

transfer vehicles (shown in Figure 33) and

"common" gas-cooled reactor technology for both

the bimodal engines and surface power reactor

system helps reduce costs further. With its

integrated power system, the bimodal core stage

also simplifies space operations and lowers mission

risk by eliminating the operational complexities of

multiple PVA "deployment / retraction" cycles

(e.g., prior to and after TMI, aerobraking and TEl

maneuvers).

With propulsive capture at Mars, the power

available in Mars orbit grows to 150 kWe per

mission--five times that of the DRM. The more

complex, higher risk aerobraking and capture

maneuver is also replaced by a simpler atmospheric

reentry using a "standardized," lower mass

"aerodescent" shell. The introduction of TransHab

and LANTR affords further mission flexibility and

downstream growth capability. The BNTR / TransHab combination provides options for reusing the

ERV and shortening its mission duration by having

the crew travel to and from Mars on the same

bimodal transfer vehicle. The addition of LANTR

engines enhances the performance of "volumelimited" vehicles by increasing their bulk propellant

density. Using bimodal LANTR and TransHab,

Phobos rendezvous and landing options can be

added to the current DRM.

If water is discovered on Phobos and its

extraction for return propellant proves feasible, then

Phobos could become an important staging point

for the future exploration and development of Mars.

A Phobos station and propellant depot would

provide reusable LANTR-powered Mars transfer

vehicles with their return propellant allowing them to

shorten trip times or transport more high value cargo

to Mars instead of bulk propellant. Reusable biconicshaped LANTR-powered ascent / descent vehicles,

operating from specially prepared sites on Mars,

would ferry modular payload elements to and from

the surface. Should Phobos be dry, they would

also resupply orbiting transfer vehicles with propellants needed to reach refueling depots in the asteroid

belt (see Figure 34). From there, the LANTRpowered transfer vehicles could continue on to the

"water rich" moons of the Jovian system, providing

a reliable foundation for the development and

eventual human settlement of the Solar System.

NASA/TM-- 1998- 208834/REV 1 42
Earth . .+_:_,::.::._:._:._:.::.:_,:.:x._x.:_,:.:.:_,_,:+:.:.:.:.:.,.:.:.:.:.:+:,_,_:.:.:,_.:.:,_:.:,_:.:,_:.;:.:_.:,_:,_:,;:,_:,_:,_:.. .+::._:.::.:x._x:_':::_:_'_'_'_:::'_::'_:.;:.;:,_::.;.

............. Moon .................................................. i::::_.1%_:_.. M_rs ......................... Asteroids:_! ............__ '::_:

__ _ _ " %:Callisto

LUN:ox

_;'_ /Polar Ice +i !!i; H20 ........ _"_ _

LH2 and LOX _ _

H20 Phobos_ Deimos Jupiter _!ii!_ _ _

Europa Ganymede

Figure 34.--Human Expansion Possibilities with LANTR Transfer Vehicles.

REFERENCES 7.

1. The Space Science Enterprise Strateqic Plan--

Oriqin, Evolution and Destiny of the Cosmos

and Life, National Aeronautics and Space

Administration (November 1997).

2. Beyond Earth's Boundaries - Human Exploration of the Solar System in the 21st Century,

Office of Exploration Annual Report to the

Administrator, National Aeronautics and Space

Administration (1988).

3. S.K. Borowski, "Impact of Solid Core Nuclear

Thermal Rocket (SC/NTR) Propulsion on Human Expeditions to Phobos / Mars," in

Exploration Studies Technical Report--FY 1988

Status, Office of Exploration, NASA Technical

Memorandum 4075, (December 1988),

pp. 5-11 to 5-17.

4. A. Cohen, et al., Report of the 90-Day Study on

Human Exploration of the Moon and Mars,

National Aeronautics and Space Administration

(November 1989).

5. S.K. Borowski, "An Evolutionary Lunar-to- Mars

Space Transportation System Using Modular

NTR/ Stage Components," AIAA-91-3573,

American Institute of Aeronautucs and Astronautics (Jan. 1991 ).

6. America at the Threshold - America's Space

Exploration Initiative, Report of the Synthesis

Group, Available from the Superintendent of

Documents, U. S. Government Printing Office,

Washington, DC 20402 (June 1991).

8.

9.

10.

J.K. Soldner and B.K. Joosten, "Mars Trajectory

Options for the Space Exploration Initiative,"

AAS-91-438, ASS/AIAA Astrodynamics Specialist Conference (Aug. 1991).

B.K. Joosten, B.G. Drake, D.B. Weaver and

J.K. Soldner, "Mission Design Strategies for the

Human Exploration of Mars," IAF-91-336, 42nd

Congress of the International Astronautical

Federation (Oct. 1991 ).

R.M. Zubrin, D. Baker, and O. Gwynne, "Mars

Direct: A Simple, Robust, and Cost-Effective

Architecture for the Space Exploration Initiative," AIAA-91-0326, American Institute of

Aeronautics and Astronautics (Jan. 1991 ).

S.K. Borowski, et al., "Nuclear Thermal Rocket/

Vehicle Design Options for Future NASA

Missions to the Moon and Mars," AIAA-93-

4170, American Institute of Aeronautics and

Astronautics (Sept. 1993) and NASA Technical

Memorandum 107071 (Sept. 1995).

11. Human Exploration of Mars: The Reference

12.

Mission of the NASA Mars Exploration Study

Team, S. J. Hoffman and D. I. Kaplan, eds.,

NASA Special Publication 6107 (July 1997).

Reference Mission Version 3.0 Addendum to

the Human Exploration of Mars: The Reference

Mission of the NASA Mars Exploration Study

Team., B.G. Drake, ed., Exploration Office

Document EX13-98-036 (June 1998).

NASA/TM-- 1998-208834/REV 1 43
13. D.R. Koeing, "Experience Gained from the

Space Nuclear Rocket Program (Rover),"

LA-10062-H, Los Alamos National Laboratory

(May 1986).

14. J.S. Clark, S.K. Borowski, R.J. Sefcik and

T.J. Miller, "A Comparison of Technology

Development Costs and Schedule for Nuclear

Thermal Rockets for Missions to Mars," AIAA93-2263, American Institute of Aeronautics and

Astronautics (June 1993).

15. J.S. Clark, M.C. Mcllwain, V. Smetanikov,

E.K. D'Yakov, and V.A. Pavshook, "US/CIS Eye

Joint Nuclear Rocket Venture," Aerospace

America, Vol. 31, (July 1993).

16. D.W. Culver, et al., "Development of Life Prediction Capabilities for Liquid Propellant Rocket

Engines," Task No. 8 (NTRE Extended Life

Feasibility Assessment), Aerojet Propulsion

Division Final Reports under NASA Contract

NAS3-25883 (Oct. 1992 and July 1993).

17. D.W. Culver, V. Kolganov, and R. Rochow,

"Low Thrust, Deep Throttling, US / CIS Integrated NTRE," 11th Symposium on Space

Nuclear Power Systems, Albuquerque, New

Mexico, (January 9-13, 1994).

18. S.K. Borowski, et al., "A Revolutionary Lunar

Space Transportation System Architecture Using Extraterrestrial LOX-Augmented NTR

Propulsion," AIAA-94-3343, American Institute

of Aeronautics and Astronautics (June 1994)

and NASA Technical Memorandum 106726

(August 1994).

19. S.K. Borowski, D.W. Culver and M.J. Bulman,

"Human Exploration and Settlement of the Moon

Using LUNOX-Augmented NTR Propulsion,"

12th Symposium on Space Nuclear Power and

Propulsion Systems, Albuquerque, New Mexico,

(January 8-12, 1995) and NASA Technical

Memorandum 107093 (October 1995).

20.

21.

22.

23.

24.

25.

26.

D.W. Plachta, personal communications, NASA

Lewis Research Center (1998).

R.H. Knoll, R.J. Stochl, and R. Sanabria,

"A Review of Candidate Multilayer Insulation

Systems for Potential Use on Wet-Launched

LH 2 Tankage for the Space Exploration Initiative Lunar Mission," AIAA-91-2176, American

Institute of Aeronautics and Astronautics

(June 1991 ).

R.M. Zubrin, "The Use of Low Power Dual Mode

Nuclear Thermal Rocket Engines to Support

Space Exploration Missions," AIAA-91-3406,

American Institute of Aeronautics and Astronautics (Sept. 1991 ).

M.L. Stancati, personal communications,

Science Applications International Corporation

(1998).

M.L. Stancati and J.T. Collins, "Mission Design

Considerations for Nuclear Risk Mitigation," Proc.

Nuclear Propulsion Technical Interchanqe Meetincl, R.R. Corban, ed., NASA Conference

Publication 10116, October 20-23, 1992, Vol.1,

pp. 358-365.

B. O'Leary, "Rationales for Early Human

Missions to Phobos and Deimos,"in Lunar Bases

and Space Activities of the 21st Century,

Lunar and Planetary Institute, Houston, 1985,

pp. 801-808.

P. Moore, G. Hunt, I. Nicolson, and P. Cattermole,

The Atlas of the Solar System, Crescent Books,

New York, 1990, pp. 240-241.

NASA/TM-- 1998- 208834/REV 1 44
APPENDIX

Table A-I. Earth Return Vehicle Pa,,Ioad Mass (kg)

Habitat Element

Life Support System

Crew Accom. + Consumables

Health Care

EVA equipment

Comm/info Management

30 kW PVA power system

Thermal Control system

Structure

Science equipment

Spares

Total Cargo Mas_

TEl stage dry mass

Propellant mass

Earth return RCS propellant

Aerobrake

Tota Pay oad Mass

26581

4661

12058

0

243

32O

3249

55O

5500

6OO

1924

29105

48O6

28866

1115

10180

74072

Table A-2. Cargo Lander Payload Mass (kg)

Earth Entry/Mars Ascent Capsule

Ascent Stage Dry Mass

ISRU plant

Hydrogen feedstock

PVA keep-alive power system

160 kW nuclear power plant

1.0 km power cables, PMAD

Communication system

Pressurized Rover

Inflatable Laboratory Module

15 kWe DIPS cart

Unpressurized Rover

3 teleoperable science rovers

Water storage tank

Science equipment

Total Cargo Mas_

Vehicle structure

Terminal propulsion system

Total Landed Mass

Propellant

Forward aeroshell

Parachutes and mechanisms

Total Payload Mass

4829

4069

3941

5420

825

11425

837

32O

0

3100

1500

55O

1500

150

1770

40236

3186

1018

44440

10985

9918

7OO

66043

Table A-3. Piloted Hab Lander Payload Mass (kg)

Habitat element 2

Life Support System

Health Care

Crew Accomodations

EVA equipment

Comm/info management

Power

Thermal

Structure

Science

Spares

Crew

3 kW PVA keep-alive power

Unpressurized rover

EVA consumables

EVA suits

Total Cargo Mas_

Vehicle structure

Terminal propulsion system

Total Landed Mass

Propellant

Forward aeroshell

Parachutes and mechanisms

Total Payload Mass

28505

4661

0

12058

243

32O

3249

55O

5500

0

1924

5OO

0

550

446

940

30941

3186

1018

35145

11381

13580

7OO

60806

NASA/TM-- 1998-208834/REV 1 45
Table A-4. "Three Mission" IMLEO Summary for "All BNTR" Option

("2 - Perigee Burn" Earth Departure Scenario/Transit Times: 220 (OB) & 180 (IB) Days)

(IMLEO < 178 t/2-88 t Magnum/Shuttle C HLLVs)

PayloadNehicle

Propulsion/Isp

Earth Return

Vehicle

Payload

Ascent Stage

LOX/CH4

Isp = 379 s

(O/F = 3.5:1)

Descent Stage

LOX/CH 4

Isp = 379 s

(O/F = 3.5:1)

Common

NTR Vehicles

w/Modular

Components

CIS w/LH2

Isp = 940 - 955 s

RCS

NTO/MMH

Isp = 320 s

Element

Masses (t)

2011

ERV

Mission

Crew Hab Module

2011

Cargo Lander

Mission

18.15

Spare ECRV

Contingency Consumables 7.31

Crew (6) & Suits

MAV Crew Cab/ECRV 4.83

Ascent Stage 4.06

Propellant* 38.40

Surface Payload 31.34

2014

Piloted Lander

Mission

1.44

26.54

Descent Stage 4.20 4.20

Aerodescent Shell + 8.23 7.94

Parachutes 0.70 0.70

Propellant** 8.91 7.92

Total Payload Mass 62.27 25.46 48.74

CIS Engines (#) 7.67(3) 7.67(3) 7.67(3)

F(klbf) per engine/Isp(s) 14.76/955 15/940 14.76/955

Radiation Shields (#) 3.24(3) 3.24(3)

"In-Line" TMI LH 2 8.25 8.52 8.25

Tank & Structure

TMI "Core" Stage 11.77 11.77

Tank & Structure

TMI/MOC/TEI "Core" Stage 11.77

Tank & Structure

Brayton Power 1.55 1.55 1.55

System (@ 50 kWe)

0.60 LH2 Refrigeration

System***

Avionics & Aux. Power

0.34

1.69 1.69

0.60

1.69

Propellant .... 68.35 62.35 77.54

Propulsion & Tankage 0.52 0.55 0.52

Propellant 1.62 2.10 1.55

Total NTR Vehicle Mass 102.02 99.78 114.38

Total IMLEO 164.29 125.24 163.12

* Produced at Mars usin( "in-situ" resources

** Assumes parachutes and 632 m/s descent AV

*** Cooling capacity of "core"/"in-line" tanks @ -75/46 Wt, respectively

.... Contains boiloff, cooldown, "tank trapped" residual and disposal LH2 also

+ Using ARC Triconic aerobrake mass estimation formula with V.=4.5 km/s

NASA/TM-- 1998- 208834/REV 1 46
Table A-5. "Three Mission" IMLEO Summary for "Alternative Mission Profile"

("Single Burn" Earth Departure Scenario/Transit times: 210 (OB) & 180 (IB) days)

(IMLEO < 160 t/2-80 t Magnum/Shuttle C HLLVs)

2011 2011 2014

PayloadNehicle Element Cargo "Unpiloted" Piloted

Propulsion/Isp Masses (t) Lander Hab Lander Mission

Earth-Mars "TransHab" Module 14.96

Transit Vehicle Crew (6) & Suits 1.44

Payload Extra Consumables 0.40

Contingency Comsumables 6.69

Ascent Stage MAV Crew Cab/ECRV 4.83

LOX/CH4 Ascent Stage 4.10

Isp = 379 s Propellant* 38.40

(O/F = 3.5:1)

Surface Payload 24.42 25.37

Descent Stage Descent Stage 4.10 4.10

LOX/CH4 Aerodescent Shell + 8.05 7.90

Isp = 379 s Parachutes 0.70 0.70

(O/F = 3.5:1) Propellant** 7.53 7.76

Total Payload Mass 53.73 45.82 23.49

CIS Engines (#) 7.67(3) 7.67(3) 7.67(3)

Common F(klbf) per engine/Isp (s) 14.76/955 14.76/955 14.76/955

NTR Vehicles Radiation Shields (#) 3.24(3)

"In-Line" TMI LH 2 4.90 4.90 8.52

w/Modular Tank & Structure

TMI "Core" Stage 11.77 11.77 Components Tank & Structure

TMI/MOC/TEI "Core" Stage 11.77

LH 2 NTR Tank & Structure

Isp = 955 s Brayton Power 1.55 1.55 1.55

System (@ 50 kWe)

LH2 Refrigeration 0.55 0.34 0.34

System***

Avionics & Aux. Power 1.69 1.69 1.69

LH2 Propellant .... 64.34 57.31 78.71

RCS Propulsion & Tankage 0.51 0.50 0.55

NTO/MMH Propellant 1.44 1.30 2.11

Isp = 320 s

Total NTR Vehicle Mass 94.42 87.03 116.15

Total IMLEO 148.15 132.85 139.64

* Produced at Mars using "in-situ" resources

** Assumes parachutes and 632 m/s descent AV

*** Cooling capacity of "core"/"in-line" tank @ -75 and 27 Wt, respectively

.... Contains boiloff, cooldown, "tank trapped" residual and disposal LH 2 also

+ Using ARC Triconic aerobrake mass estimation formula with V_ = 4.5 km/s

NASA/TM-- 1998-208834/REV 1 47
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1. AGENCY USE ONLY (Leave blank) 2. REPORT DATE 3. REPORT TYPE AND DATES COVERED

December 2002 Technical Memorandum

4. TITLE AND SUBTITLE 5. FUNDING NUMBERS

Vehicle and Mission Design Options for the Human Exploration

of Mars/Phobos Using "Bimodal" NTR and LANTR Propulsion

6. AUTHOR(S)

Stanley K. Borowski, Leonard A. Dudzinski, and Melissa L. McGuire

7. PERFORMING ORGANIZATION NAME(S) AND ADDRESS(ES)

National Aeronautics and Space Administration

John H. Glenn Research Center at Lewis Field

Cleveland, Ohio 44135-3191

9. SPONSORING/MONITORING AGENCY NAME(S) AND ADDRESS(ES)

National Aeronautics and Space Administration

Washington, DC 20546-0001

WU-953-20-0C-00

8. PERFORMING ORGANIZATION

REPORT NUMBER

E- 11445-1

10. SPONSORING/MONITORING

AGENCY REPORT NUMBER

NASA TM--1998-208834-REV1

AIAA-98-3883

11. SUPPLEMENTARY NOTES

Prepared for the 34th Joint Propulsion Conference cosponsored by the AIAA, ASME, SAE, and ASEE, Cleveland, Ohio, July 13-15,

1998. Stanley K. Borowski and Leonard A. Dudzinski, NASA Glenn Research Center; Melissa L. McGuire, Analex Corporation,

3001 Aerospace Parkway, Brook Park, Ohio 44142 (work funded under NAS3-27186). Responsible person, Stanley K. Borowski,

organization code 6510, 216-977-7091.

12a. DISTRIBUTION/AVAILABILITY STATEMENT 12b. DISTRIBUTION CODE

Unclassified -Unlimited

Subject Categories: 12, 15, 16, and 20 Distribution: Nonstandard

Available electronically at httlg_://gltrs_grc.nasg_,gov

This publication is available from the NASA Center for AeroSpace Information, 301-621-0390.

13. ABSTRACT (Maximum 200 words)

The nuclear daermal rocket (NTR) is one ofdae leading propulsion options for future human missions to Mars because of its high specific impulse (lsp-850-1000 s)

capability and its at*xactive engine tlrrust-to-weight ratio (-3-10). To stay wi/Jain the available mass and payload volume limits of a "Magnum" heavy lift vehicle, a high

performance propulsion system is required for trans-Mars injection (TMI). An expendable TMI stage, powered by dtree 15 daousand pounds force (klbf) NTR engines

is currently under consideration by NASA for its Design Reference Mission (DRM). However, because of the miniscule bttrnup of enriched nranium-235 during the

Earth departure phase @10 grams out of 33 kilograms in each NTR core), disposal of the TMI stage and its engines after a single use is a costly and inefficient use of

this high performance stage. By reconfigrmng the engines for bo/Ja propulsive tkntst and modest power generation (referred to as "bimodal" operation), a robust, multiple

burn, "power-rich" stage wi/Ja propulsive Mars capture and reuse capability is possible. A family of modular "bimodar' NTR (BNTR) vehicles are described which utilize a common "core" stage powered by tlrree 15 klbf BNTRs/Nat produce 50 kWe of total electrical power for crew" life support, an active refrigeration / reliquification

system for long term, zero-boiloff liquid hydrogen (LH2) storage, and high data rate communications. An innovative, spine-like "saddle truss" design connects dae core

stage and payload element and is open undernea/Ja to allow supplemental "in-line" propellant tanks and contingency crew" consumables to be easily jettisoned to improve

vehicle performance. A "modified" DRM using BNTR transfer vehicles requires fewer transportation system elements, reduces IMLEO and mission risk, and simplifies

space operations. By taking the next logical step_se of the BNTR for propulsive capture of all payload elements into Mars orbit the power available in Mars orbit

grows to 150 kWe compared to 30 kWe for the DRM. Propulsive capture also eliminates the complex, higher risk aerobraking and capture maneuver which is replaced

by a simpler reentxy using a standardized, lower mass "aerodescent" shell. The attractiveness of the "all BNTR" option is further increased by the substitution of the

lightweight, inflatable "TransHab" module in place of the heavier, hard-shell hab module. Use of TransHab intxoduces the potential for propulsive recovery and reuse

of the BNTR / Earth return vehicle (ERV). It also allows the crew" to travel to and from Mars on the same BNTR transfer vehicle/Jaereby cutting tJaedtmation of tJae ERV

mission in hal_from -4.7 to 2.5 years. Finally, for difficult Mars options, such as Phobos rendezvous and sample return missions, volume (not mass) cous/xaints limit

dae performance of dae "all LH2" BNTR stage. The use of "LOX-angmented" NTR (LANTR) engines, operating at a modest oxygen-to-hydrogen mixture ratio (MR)

of 0.5, helps to increase "bulk" propellant density and total tlrrust during the TMI burn. On all subsequent burns, the bimodal LANTR engines operate on LH 2 only

(MR=0) to maximize vehicle performance while staying widain dae mass limits of two Magnum launches.

14. SUBJECT TERMS

Nuclear thermal rocket; NTR; "LOX-augmented" NTR; LANTR; Bimodal;

In-situ resource utilization; High thrust; Nuclear propulsion; Mars; Phobos; Spacecraft

17. SECURITY CLASSIFICATION 18. SECURITY CLASSIFICATION 19. SECURITY CLASSIFICATION

OF REPORT OF THIS PAGE OF ABSTRACT

Unclassified Unclassified Unclassified

NSN 7540-01-280-5500

15. NUMBER OF PAGES

53

16. PRICE CODE

20. LIMITATION OF ABSTRACT

Standard Form 298 (Rev. 2-89)

Prescribed by ANSI Std. Z39-18

298-102