(PDF) Mars 2033 Human Flyby Mission

73st International Astronautical Congress (IAC) – 18-22 September 2022.

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IAC-22,B3,8,x70674

Mars 2033 Human Flyby Mission

Benjamin Donahuea, Matt Dugganb

a Boeing Exploration Launch Systems, Huntsville, Alabama, USA, benjamin.b.donahue@boeing.com

b Boeing Space Systems, Houston, Texas, USA, matt.b.duggan@boeing.com

* Corresponding Author

Abstract

A 2033 Mars Flyby mission is described and the characteristics of the trajectory, in-space stages, crew habitat and

Earth Entry Capsule are described. An Earth Mars alignment, that allows for a free return trajectory at Mars, occurs

every 15 years. This means that no propulsive maneuver is required at Mars to affect a return to Earth. The vehicle

swings by Mars and is thereafter on a path to intercept the Earth. This greatly reduces the energy requirements for the

transfer stages, and allows for lighter stages and fewer launches than which would be required for a stopover

mission. With a trip time of 530 days, this 2033 crewed flyby mission could serve as a precursor to a later Mars

surface mission. Four elements are required; a crew habitat, Earth entry capsule, in-space stage and an Earth

departure stage.

Keywords: (NASA, exploration, space launch system)

Acronyms /Abbreviations

Astronomical Unit (AU)

Beyond Earth Orbit (BEO)

Delta-Velocity (dV)

Reusable Solid Rocket Motor (RSRM)

Space Launch System (SLS)

Injection energy (C3)

Exploration Upper Stage (EUS)

Beyond Earth Orbit (BEO)

Oxygen (O2)

Hydrogen (H2)

Trans-Mars Injection (TMI)

Earth Entry Capsule (EEC)

Thermal Protection System (TPS)

Deep Space Burn (DSB)

Earth Moon Lagrange Point 2 (EM-L2)

Near Rectilinear Halo Orbit (NRHO)

Space Launch System (SLS)

1. Introduction

This report has a three-fold objective: 1) To describe

a crewed Mars flyby mission departing in 2033; 2) to

describe several key options involving in-space

stages, trajectory, and the launch manifesting of the

elements making up the Mars transfer system, and 3)

to describe the launch vehicle for this mission, the

NASA Space Launch System (SLS).

2. SLS: A family of Evolved Launchers

The SLS system has been developed with planned

upgrades to increase mass and volume lift capability

that could be used to enhance or enable a variety of high

energy missions. The initial Block-1 configuration

delivers an 85-metric ton (mt) class (to LEO) capability.

Block upgrades are already underway to bring online a

Block 1B, 105 mt class (to LEO) capable configuration

in 2026, and a Block 2, 130 mt class (to LEO) capability

version later in the decade. The evolutionary path of the

SLS is shown in Fig. 1 and 2.

2.1 The Initial Block 1 Variant of the SLS

The Block-1 variant is used for the first integrated

flight of the full SLS system, known as Artemis I.

Block-1 will also be used for the second flight, which

will be the first human launch for SLS with the Orion

crew vehicle, and for the third flight, which will mark

the return of humans to the lunar surface.

2.2 The Evolved Block 1B Variant of the SLS

The initial Block-1 configuration evolves into a

higher performing version with the addition of a new,

larger upper stage, the Exploration Upper Stage (EUS).

The EUS will contain 114 mt (250,000 lb) of propellant

and will be powered by four RL10‐C engines. With

increased propellant and engine thrust, the EUS

significantly increases the capability of SLS compared

to Block-1. This more powerful configuration is known

as Block-1B (Fig. 1). In its cargo configuration, the

Block 1B adds a much larger 8.4-meter fairing, greatly

increasing the volume available for payload. The

crewed variant will utilize the Orion capsule to transport

astronauts. The more capable Block-1B vehicle with

EUS will first fly in 2026 or 2027.

2.3 The Performance Enhanced Block 2 SLS

The current SLS boosters are improvements to the

RSRM designed and built for the Space Shuttle. When

applied to the SLS architecture, increased lift capacity is

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achieved by removing systems necessary for booster

recovery and reuse that were employed during the

Shuttle program. Eight flight sets of shuttle booster

hardware (steel cases, structures, etc.) are available for

use before parts and processes need to be redesigned,

requalified, and rebuilt. NASA is currently studying

Booster Obsolescence and Life Extension (BOLE) of

the SLS boosters [1]. The study is looking to replace

obsolete materials and to include technologies to

improve the manufacture and performance of the

boosters, with upgrades to the propellant and liner

system. The transition from Block-1B to Block-2 by

replacing the current boosters with BOLE boosters is

shown in Fig. 2.

3. 2033 Mars Flyby Opportunity

Earth-Mars mission opportunities are linked to the dates

of alignment of the two bodies when Mars and the sun

are in opposite directions as viewed by an observer on

Earth. This alignment is termed opposition alignment,

or simply opposition. Alignments for which Mars and

the sun are in the same direction are termed superior

conjunction, or simply conjunction. Occurrences of

either type of alignment repeat at 26-month intervals

(the synodic period of the Earth-Mars system), on

average, with variations caused by Mars orbit

eccentricity. Successive oppositions occur at longitudes

that progress around the sun, completing a full

revolution after seven oppositions and fifteen years.

Because of the eccentricity of Mars, this produces a 15-

year cycle of varying performance requirements of

successive missions such that some opportunities in the

cycle are considered better than others.

Low energy transfers in either direction begin prior

to an alignment date and end after the alignment date.

The alignment occurs at about the midpoint of the

transit trajectory arc. Because the spacecraft travels

slower as the distance from the sun increases, the time

between Earth departure and the alignment date is less

than that from the alignment date to Mars arrival. For a

typical transfer of seven months, launch would occur

about 2.5 months prior to the alignment; on a return

flight, Mars departure would occur about 4.5 months

prior to an alignment. Because Mars arrival occurs after

the alignment of the outbound transfer, a low energy

return trajectory back to the Earth must be linked to the

opposition following that of the outbound trajectory.

Combining these two trajectories leads to typical

mission durations of about 31 months (~940 days) and

stay times of 17 months (~515 days). This is the classic

conjunction class mission, so called because

conjunction occurs about midway during the stopover

period. For manned missions to Mars, there has long

been interest in shorter mission durations and stay times

than is afforded with conjunction class missions. These

shorter missions are known as opposition class missions

because both the outbound and return legs are linked to

the same opposition. Typically, an opposition class

Mars mission profile consists of a low energy outbound

trajectory, followed with a short duration stopover

period of less than 60 days, and ends with a higher

energy, longer duration return trajectory that passes

through perihelion near the orbit of Venus before

returning to Earth. Performance requirements for the

mission increase substantially with stay time, so the

goal is to keep the stay time as short as possible,

consistent with science objectives. The reverse sequence

(Venus swingby on the outbound leg) of this mission is

also possible, but that has received less interest because

the highest delta-velocity (dV) requirement then occurs

at launch rather than at the end of the mission where the

vehicle mass is much less.

Abort studies of early opposition class missions

identified the possibility of a free return trajectory that

could be employed in the event of a propulsion system

failure on the outbound leg of the mission. Such a

mission is simply a variation of the opposition class

mission profile that eliminates the stopover and, instead,

employs a gravity assist of Mars, using no propulsion

following Earth departure. Analysis of this concept has

determined that it offers potentially viable solutions in

only two of the seven oppositions of the 15-year cycle

described above. For the other five oppositions of the

cycle, the Earth Departure launch Δv is excessive, or the

required passage distance at Mars falls below the

surface radius. The most recent potentially viable

solutions were associated with the opposition years of

2016 and 2018; the earliest future opportunities of

interest occur for opposition years 2031 and 2033. For

both of these pairs, the latter oppositions (2018 and

2033) yield the lesser propulsion requirements. The

2033 opposition serves as the basis for the proposed

inaugural manned Mars mission and is the opposition

year chosen for this analysis.

3.1 Mission Parameters

Although the idea of a free return trajectory is

appealing, the mission contains some elements that

make it impossible to implement directly with today’s

technology. First, a true free return trajectory implies

that the crew would return to Earth using direct entry.

However, a characteristic of Mars free return

trajectories is high Earth arrival speeds that result in

atmospheric entry speeds greater than 14 km/s. The

Orion crew capsule can accommodate entry speeds up

to 12.5 km/s, provided atmospheric braking turns are

employed to control dynamic pressure and heating

loads. To control this entry speed, a Deep Space Burn

(DSB) propulsion maneuver, applied near the perihelion

of the return trajectory, is employed.

A second issue with free return trajectories is that all

components of the spacecraft, other than the entry

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capsule, must be discarded at the end of the mission.

Economic considerations suggest that the expensive

components of the spacecraft be retained for reuse by

recapturing the spacecraft into orbit. Doing so

eliminates the need for direct entry but does require

propulsion maneuvers. Optimization of this trajectory

retains the deep space maneuver near perihelion and

adds additional maneuvers to insert the spacecraft into

final orbit.

The size of the spacecraft needed for a manned

mission dictates that its components be delivered to an

aggregation orbit, where the spacecraft is assembled.

Various aggregation orbits have been analyzed over the

last few decades, taking into consideration cost of

access, orbit stability, and suitability for lunar and more

distant target missions. Informed by these studies, the

aggregation orbit chosen for this mission is a low-

amplitude halo orbit around the Earth-Moon libration

point 2 (EM-L2). In this region of cis-lunar space, the

gravitational effects of the sun contribute substantial

reductions to enter and depart the orbit compared to

other possible orbits, such as the Near Rectilinear Halo

Orbit (NRHO), the suggested orbit for the Lunar

Gateway, if developed. The mission spacecraft will both

depart from and return to the EM-L2 orbit.

Multiple launches are required to deliver spacecraft

components to the aggregation orbit where the

spacecraft will be assembled. Each launch will place its

payload on a trajectory to the vicinity of the moon with

an energy (C3) of -1.4 km2/s2. Upon reaching the

moon, the payload will be eased into the aggregation

orbit with a maneuver requiring approximately 363 m/s

dV. Joining of the four spacecraft elements will be

performed robotically. When ready, the crew is

transported to the spacecraft and any final assembly and

checkout will be performed. The planned launch

sequence to transfer from the aggregation orbit to

prepare for the Trans-Mars injection (TMI) is that

proposed by Dunham (Ref 2) for a Mars mission,

starting from the EM-L2 orbit; EM-L2 is 61,347 km

beyond the Moon.

The sequence consists of five propulsion maneuvers

that place the spacecraft in a highly elliptic orbit with a

perigee altitude of 622 km and apogee distance of 65

Earth radii (414,579 km). The five maneuvers are

defined in Table 1 for the Habitat expended option and

Table 2 for the Habitat recovered back at Earth option.

The powered lunar swingby maneuver initiates a series

of phasing orbits that re-orient the orbit plane to

coincide with the TMI hyperbolic excess vector. These

are mission dependent and the values presented in the

table are considered representative, but realistic. The

TMI is performed at perigee of the final elliptic orbit to

deliver the C3 required for the mission. The time

between departure of the halo orbit and TMI is about 37

days. The trajectory is optimized with the MAnE™

software from SpaceFlightSolutions. The events of the

mission following TMI include the passage of Mars at

an altitude of 250 km, a deep space propulsion

maneuver near perihelion close to the orbit of Venus,

and a final capture sequence upon arrival at Earth. The

trajectory profile for the mission for the opening day of

a 21-day launch period is shown in Fig. 4 and a mission

timeline in Fig. 5. Note that the arrows appearing at

Earth departure and Earth return represent the direction

of the hyperbolic excess vectors. At Mars, the two

vectors represent the hyperbolic excess velocity vector

before and after the gravity assist event. At the DSB

event, the two vectors represent the heliocentric velocity

vectors before and after the burn. The launch C3 for this

mission is 36.35 km2/s2 and the return hyperbolic

excess speed is 3.69 km/s. For the option of Habitat

recovery at Earth, the dV at Earth return places the

spacecraft in a capture orbit with perigee altitude of 622

km and apogee distance of 65 Earth radii. The reverse

of the 5-maneuver departure sequence is used to transfer

to the original EM-L2 halo orbit.

3.2 Mission Options & Selections

Several mission options are available. The first involves

the type and number of launch vehicles used. The

second option involves the aggregation orbit for joining

of the Mars vehicle elements. The third option involves

the Earth return method: either by an Earth Entry

Capsule (EEC) or via capturing the crew habitat back

into Earth orbit, then transiting it to EM-L2. The fourth

option involves the magnitude of the DSB burn (and

dV); this impacts the Earth return arrival velocity and

the EEC TPS capability. the fifth option involves the

propellant (and engine) selection for the two in-space

stages (TMI and DSB). The sixth option involves

selection of the number of crew. The seventh option

involves the inclusion of a sunshield for the EEC during

the portions of flight with high solar flux.

The selections made for this study are as follows:

Launch vehicle types: SLS and Vulcan Heavy

 Number of Launch vehicles used: 3 or 4

 Aggregation orbit: EM-L2

 Earth Return Mode: EEC with Habitat expended as

the baseline. alternative: crew Habitat recovered

 DSB maneuver for limiting Earth arrival velocity:

dV=1,461 m/s to limit to 12.5 km/s (reference)

dV=520 m/s to limit to 13.9 km/s (alternative)

 Crew size: 2-3

 Sun Shield: Deployable shield for EEC (Fig 8)

4. Earth Entry

If an EEC is used, its Earth entry velocity limitations

must be evaluated, specifically the capability of its

Thermal Protection System (TPS) (heat shield) to

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withstand the heat loads that occur at high entry

velocities characteristic of Mars missions. Two options

are accessed. The first is a direct entry of the crew in the

EEC. Three types of EEC’s were considered; the Orion,

the Boeing CST100 and a new small Capsule

specifically designed for high velocity reentries. In the

first instance, if the Orion is modified to endure the

flight time, it could be used for the direct entry. In this

case the Habitat is not recovered at Earth For these free

return missions, the entry speeds are high; more than 14

km/s, well above Lunar entry speeds.

Since Orion uses the same ablative coatings in its

TPS as Apollo, a DSB, on the inbound a leg is required

to reduce the entry speed to 12.5 km/s (Ref 3). The DSB

burn, for the 2033 mission, is shown in Fig. 4, at a cost

of 1,465 m/s dV and occurs near periapsis at

approximately the orbit radius of Venus. In the

“Inspiration Mars” paper (Ref. 4), mission designers

recognized the difficulties of entry velocities from a

Mars mission, as they all exceed 14.2 km/s entry speeds,

well beyond the Apollo technology. They proposed

Aerocapture, using aero maneuvering to reduce the

entry environment heating.

In 2005 Putnam, Braun et al discovers there is not

much difference between aerocapture and direct entry in

terms of heating, especially using banking maneuvers

during the latter. They determine, a Lift-to-drag (L/D)

ratio of 0.3 is adequate to fly an Earth entry flight

corridor between 0.50–0.70 degrees, the required

corridor width for ballistic coefficients in the range of

300 kg/m2. The analysis that follows is based on a

direct entry with a speed of 12.5 km/s and that uses

simple banking maneuvers to reduce g-loads and

heating. The proper corridor is shown in the Fig. 3.

Describing the entry, Braun is quoted:

“The upper bound of the aerodynamic corridor,

which is achieved by flying the vehicle in a lift-down

attitude, is the shallowest trajectory that the vehicle can

fly while still achieving the proper energy decrement. In

this manner, the vehicle stays in the atmosphere the

maximum time, decelerating at a nearly constant

altitude. Hence, the required energy loss, deceleration,

and heat transfer are spread out over time. The lower

corridor bound, which is attained by flying the vehicle

in a lift-up attitude, is the steepest trajectory that still

obtains the proper atmospheric exit conditions. By

following this atmospheric flight path, the vehicle

passes in and out of the atmosphere in the shortest

amount of time. Because this trajectory must lose the

same amount of energy as the lift-down transfer but in a

shorter duration, it is characterized by a higher max

deceleration and heating rate.”

Braun's corridor shows the lift–up maneuver to

reduce g load and lift-down maneuver to reduce heat

load. At some point (75,000 meters altitude), we can

approximate the transfer through the corridor as an

equilibrium glide model. This approximation allows an

analytic approach that can be applied to Braun’s

numerical analysis. These relationships provide the

proper relative velocity experienced by the Orion

Capsule. Applying this model to the parameters used in

Braun’s numerical analysis, we get the second column

in Table 2. Notice the model is conservative with

respect to the numerical results. Now, we need to adjust

for the Orion’s ballistic coefficient and nose radius.

These changes are represented in the third column.

Notice convective heating goes down while radiative

heating increases. This is due for the most part to the

larger nose radius of the Orion capsule. By performing a

deep space burn to constrain the entry speed to 12.5

km/sec, we can maintain a demonstrated, Apollo

heating environment. This is illustrated by the last entry

in the table. The total heating load is estimated to be

1350 Watts/cm2. The Apollo Avcoat 5026 material,

selected for Orion capsule, is qualified to < 2500

watts/cm2.

Another option is to replace Orion with a CST-100,

combined with a small entry capsule, specifically

configured for a high earth arrival velocity. The capsule

would utilize the Ames-developed Phenolic

Impregnated Carbon Ablator (PICA) thermal protection

material, which can be readily tailored for higher entry

speeds, negating the need for a large deep space burn.

However, there is a trade in sizing the amount of PICA

and including a ‘flyable’ corridor to reduce the heat

load. One needs a 0.5-0.75-degree corridor to be

effective. The boundary for this trade is an entry speed

of 13.9 km/s. We chose this value for our larger trades

in this paper. The last mission possibility is to bring the

Habitat back to Earth. That involves a full capture, with

a return to EM-L2. In this option, the Orion mission is

restricted to ferry the crew from EM-L2 to Earth. Orion

might be stationed at the Gateway and be summoned to

EM-L2 to retrieve the crew for a homeward journey. In

this case, the DSB Stage dV is increased to 2,115 m/s.

In addition, there is another 1,029 m/s dV required for a

capture maneuver, followed by a 269 m/s dV Oberth

maneuver to return to EM-L2. These latter maneuvers

are laid out in Table 2 and amount to 269 m/s.

5.0 Transfer Stages and Elements

Four elements are launched into EM-L2; the EEC, the

crew Habitat, the TMI stage and the DSB Stage(s).

5.1 Trans-Mars Injection (TMI) Stage

The TMI stage will be launched by a SLS

Block-2; after EUS boost to TLI, the stage

transfers to EM-L2. The Stage uses O2/H2

propellant and RL10 engines, with specific

Impulse (Isp) of 465 sec. Once at EM-L2, the

TMI stage will be mated to the other elements.

This stage is launched last in the sequence.

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5.2 Deep Space Burn (DSB) Stage

The Deep Space Burn stage is launched by an

SLS Block-2 with the Habitat. After EUS boost

to TLI, the DSB Stage/Habitat combination will

transfer to EM-L2 and later it will dock to the

EEC. The DSB Stage utilizes either space

storable NTO/MMH or soft-cryogen O2/CH4

propellant; for the former, pump-fed engines of

339 sec Isp are selected; for the latter, staged

combustion cycle engines of 385 sec Isp are

selected (Ref 6). Post-TMI, the DSB provides

all propulsive maneuvers for the remainder of

the mission until capsule separation at Earth

arrival. DSB size is dependent on the dV and

payload mass boosted to EM-L2 (burn 1) and

again at perihelion (burn 2).

5.3 Earth Entry Capsule (EEC)

The Earth Entry Capsule (EEC) is an Orion or

CST-100 launched with the Block-2 (Orion), or

Vulcan Heavy (CST-100). After EUS boost to

TLI, the EEC will either self-ferry to EM-L2 or

be boosted there by a DSB stage. For cases

using the small Earth Entry Capsule, (designed

for a 13.9 km/s arrival velocity), a CST100 is

also launched, though only for the purpose of

ferrying the crew to EM-L2, where it remains,

or returns to Gateway.

5.4 Long Duration Crew Habitat

The Crew Habitat contains all the elements,

subsystems, consumables, spares, power and

avionics necessary for life support, command

and control, telemetry, science, EVA and

medical functions. For most scenarios, the DSB

stage/ Habitat is launched first, followed by the

EEC. The TMI stage is launched last. Launch

manifest options are discussed in Section 7. On

aspect of this evaluation was to determine the

allowable mass of the Habitat given the other

parameters.

6. In-Space Maneuvers

Once aggregated at EM-L2, the Mars Flyby Vehicle

System is checked out and all systems are verified

before Earth departure. At departure, the 4-element

system is boosted to TMI velocity (C3=36.2 km2/s2;

1,947 m/s dV). Earth departure occurs on December

5th, 2032. About midway to Mars, the DSB stage

performs a mid-course correction burn (dV 12 m/s).

After a 9-month travel time, the swingby takes place

(Fig. 4), making a very close passage at an altitude 250

km above the surface on August 8th, 2033. No

propulsive maneuver is required, as the vehicle is on a

free return trajectory. Mars arrival V-infinity is 5.1

km/s. The crew might tele-operate, real time, pre-

emplaced surface assets during their pass. The inbound

trip back to Earth is 9 months; at 4 months out from

Mars the DSB Stage does another mid-course correction

burn. At 7 months out from Mars the vehicle reaches

perihelion (Venus AU=0.7, Fig 4). Here the DSB stage

fires its engines to decelerate the vehicle to reduce the

later Earth arrival velocity to 12.5 km/s. The DSB

maneuver occurs on February 4th, 2034. Once the burn

is completed, the DSB stage is retained for its RCS and

Earth arrival aim point maneuver capabilities. Just

before the DSB, collected waste is jettisoned to lighten

the Habitat.

Later, at Earth arrival (June 2034), the crew in the

EEC separates from the DSB Stage-Habitat and

reenters. The EEC follows its entry corridor and ends

the mission in a splashdown. Options relating to Earth

return were investigated. By providing additional

propellant to the DSB stage, more dV can be applied at

perihelion to achieve lower arrival velocities if

necessary. Also, by carrying a capsule specifically

designed for higher Earth arrival velocities (13.9 km/s)

less DSB propellant need be carried. Reducing DSB

mass will allow a heavier Habitat to be carried. Arrival

velocity capability is in part a function of the Capsules

ballistic coefficient and the thickness of its heat shield.

A smaller diameter, lighter weight Capsule with a more

robust TPS (certified to a higher arrival velocity), would

allow the downsizing of the DSB stage.

For a sufficiently capable EEC (>14.5 km/s

velocity), the DSB maneuver might be eliminated

altogether. Trading the propellant load of the DSB stage

against a new design small Capsule specifically

designed for high speed Earth entry is in work. Also,

increasing the thickness of the Orion’s and the CST-

100’s TPS (such that either of these could withstand

entry velocities>12.5 km/s) are also under investigation.

Presently, the CST100’s heat shield masses about 600

kg; increasing its thickness so as to double its mass

might be one potential approach to increasing its arrival

speed limit and thus reducing required DSB propellant

and stage mass. Increasing the thickness of the Orion’s

TPS is also an option. (The Orion at 27 mt fully fueled,

is significantly heavier than a CST-100).

One must also consider the total weight that the

Orion’s parachutes can decelerate. For staying within

the parachute limit, it might be possible to ‘de-content’

the Orion so that an increase in its heat shield mass is

off-set, and the new total mass would not be too heavy

for its parachutes. To summarize, we have five EEC

options: the 1) Orion, the 2) CST-100, 3) a small

capsule specifically for a high velocity, and modified, 4)

lower mass versions of Orion and 5) CST-100. Each of

the two existing Capsules have a Service Module (SM);

each of these might be flown with a reduced propellant

load to decrease the mass injected to Mars. For

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example, the Orion’s SM nominally carries 8.6 mt of

propellant; off-loading 50% would reduce Orion to 22.4

mt while retaining full functionality. (There are

scenarios in which the entire 8.6 mt prop load would not

be required). The propellant savings could then be

allocated to another element, such as the Habitat.

For options that involve the recovery of the Habitat

at Earth, an EEC is not carried to Mars. In this option

the crew is ferried to EM-L2 in the EEC, but the EEC

remains in EM-L2 or returns to Gateway. The DSB

Stage would carry additional propellant for the added

dV needed to capture the Habitat back into Earth orbit,

and transition to EM-L2 for rendezvous with the

awaiting EEC. There the Hab is refurbished for reuse.

The EEC ferries the crew to a landing.

7. Launch Manifests

Over a dozen launch manifests have been examined; six

examples are given below. What is to be shown are the

launches necessary to emplace the Hab, EEC and

transfer stages into the aggregation orbit. The launch

vehicles (LV) boost their payloads to TLI velocity, and

as mentioned, from there the transfer stages (or EEC)

boost themselves and their payloads to the aggregation

orbit. (SLS launch margin for all cases is 2.6 mt). The

dV needed from post-TLI travel to EM-L2 is 313 m/s,

plus 55 m/s dV for rendezvous.

What is to be determined in this analysis is the

Habitat mass that can be taken to Mars. Transfer stage

mass is a function of dV, propulsion efficiency, mission

duration, and is limited by the Launch Vehicles payload

capability to TLI. (the in-space stage can mass no more

than the total launch to TLI capability minus the mass of

any all elements launched with the stage, minus the

mass of the payload attach fitting). For each mission set

an EEC is selected (Orion, CST100 or small capsule).

The first conclusion of the Launch Manifest analysis

is that the 2033 crewed Mars flyby mission can be done

with 3 launches; a Habitat can be carried that is

sufficient for the 531-day mission. Adding a fourth

launch allows a heavier Habitat. The mass increase

could come as a physically larger module, and/or

additional redundancy, spares, and consumables, or on-

board science.

7.1 Mission Sets

Six mission sets are illustrated in Fig’s 10-15. The first

three sets are three launch solutions. The first of these 3-

launch sets, CASE-1, uses an Orion EEC. Three SLS

Block-2 launchers are utilized (Fig. 10). The second 3-

launch scenario, CASE-2 (Fig. 11), uses a CST100

EEC. Two Bk-2s are used along with a Vulcan Heavy

launcher for the CST100. The third 3-launch set, CASE

3 (Fig. 12) utilizes a small capsule for EEC, capable of

an Earth arrival velocity of 13.9 km/s; this solution also

uses a Vulcan Heavy for crew delivery to EM-L2 via a

CST100 (which does not travel to Mars).

7.2 Three Launch Mars Flyby Scenarios

CASE 1: Fig. 10

Orion Capsule to Mars, Earth Reentry

Habitat Expended

TMI and DSB Stages

Given three SLS Block-2 launches, an EM-L2

aggregation orbit, the Orion, a 12.5 km/s Earth arrival

limit, and selecting LO2/LH2 and NTO/MMH

propulsion for the TMI and DSB stages respectfully,

allows an 18.3 mt Habitat to be sent to Mars. DSB-1

Stage mass is 26.1 mt (it is launched with the Habitat as

the first launch). The second launch is a Block-2

carrying the Orion. Since the Orion does not need its

full propellant load for this mission, its SM is off-loaded

to 60% of its capacity. This off-loaded Orion (23.2 mt)

is launched with the DSB-2 Stage (17.7 mt) and the

Universal Spacecraft Adaptor (USA) (4.1 mt). TMI

Stage mass is 44.4 mt.

The DSB stage provides 1,645 m/s dV at perihelion

to limit the arrival velocity to 12.5 km/s. The Mars

Flyby vehicle (DSB-1 Stage/ DSB-2 Stage/ Habitat/

Orion), after an 8-month journey, passes by Mars at an

altitude of 250 km. The Orion serves as a contiguous,

backup crew command center to the primary Habitat,

offering dissimilar redundancy in systems; including

ECLSS, command/control, vehicle health monitoring,

RCS, power, telemetry and others. A Sun shield (Fig. 8)

is deployed over the Orion for perihelion passage,

where the Solar flux is three times that at Earth AU.

CASE 2: Fig. 11

CST100 to Mars, Earth Reentry

Habitat (main & supplemental) Expended

TMI and DSB Stage

Two SLS launches and a single Vulcan Heavy launch,

allow a 19.5 mt Habitat to be sent to Mars. This Habitat

is launched in two pieces; a module launched with the

DSB stage, and a smaller module launched with the

TMI stage. DSB stage mass is 28.7 mt; TMI Stage 40.6

mt. CST100 is 13.8 mt (offloaded). The DSB Stage

provides 1,645 m/s dV at perihelion so that the arrival

velocity at Earth is 12.5 km/s.

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CASE 3: Fig. 12

Small Earth Reentry Capsule to Mars

CST100 to Aggregation Orbit

Habitat Expended

TMI and DSB Stage

For this case, allowable Habitat mass is 23.2 mt; it is the

first launched on the SLS along with a 21.2 mt DSB

Stage. The CST100 (14.5 mt) is launched on a Vulcan

Heavy; it remains in EM-L2 or returns to Gateway. The

TMI Stage (34.4 mt) is launched last with a 10.0 mt

Small EEC, which reenters at 13.9 km/s (DSB dV for

this case is 520 m/s, a significant reduction from the

1,645 m/s needed for the slower 12.5 km/s Earth entry).

This option provides the highest Habitat mass of the

three launch scenarios detailed here.

7.3 Four Launch Mars Flyby Scenarios

CASE 4: Fig. 13

Orion: to Mars, Earth Reentry

Habitat Expended

TMI and Two DSB Stages

Four SLS launches allow a significantly heavier Habitat

mass (28.1 mt) for this Case-4. DSB Stage 1, 2 and 3

masses are 16.3, 44.5 and 17.5 mt; TMI Stage 44.4 mt.

A 23.4 mt Orion travels to Mars and back. The Orion’s

SM is off-loaded to 60% of its propellant capacity.

CASE 5: Fig. 14

Small Earth Entry Capsule to Mars, Earth Reentry

CST100 to Aggregation Orbit

Habitat Expended

TMI and Two DSB Stages

For this case, Habitat mass is 37.8 mt; after SLS injects

it to TLI, it is boosted to EM-L2 by a small propulsive

stage of 6.6 mt. The DSB Stage (34.1 mt) is launched

with the Small Capsule (10.0 mt). TMI Stage is 44.4 mt.

Vulcan Heavy launches a CST100 (14.5 mt) to EM-L2

(where it remains, or returns to Gateway). The DSB dV

is 520 m/s. This option provides the heaviest habitat of

all options.

CASE 6: Fig. 15

Orion to Aggregation Orbit, No Capsule to Mars

Habitat Recovered Back at Earth

TMI and Three DSB Stages

The Habitat is captured back at EM-L2 for reuse. No

capsule is carried to Mars. Four Bk-2 launches provide a

26.4 mt Habitat. This scenario requires 3 DSB Stages of

18.0, 44.5 and 15.5 mt. TMI Stage is 44.5 mt. The Orion

carries the crew to EM-L2 and there it remains, or

returns to the Gateway. This mission requires the

highest DSB dV (1,986 m/s). Earth capture (707 m/s)

and Oberth maneuvers (428 m/s) are done by the DSB-1

Stage upon Earth arrival. Back at EM-L2 the

Habitat/DSB-1 rendezvous with Orion. Hab reusability

reduces the launches needed for a future Mars mission.

8. The Impact of Missing the 2033 Opportunity

Should the vehicle elements not be ready on time for the

2033 opportunity, a 2035 opportunity is available – it

shares the chief attribute of the 2033 mission, in that no

propulsive maneuver is required at Mars passage.

Furthermore, the 2035 mission employs a close Venus

passage on the inbound leg that eliminates the need for

any propulsive maneuvers (such as a DSB) following

Earth departure. In addition, the natural arrival entry

speed at Earth is less than 12.5 km/s for all launch dates

within the launch period. This permits the direct entry of

the capsule upon return.

The trajectory diagram in given in Fig. 18. For the

opening day of the launch period, Earth departure dV

(TMI) from EM-L2 is 2,641 m/s; for any other day

during the launch period the TMI dV is less. The

passage of Venus employees a closest approach altitude

of 10,855 km. The passage at Mars is constrained to an

altitude of 250 km, for which the passage excess speed

(V-infinity) is 7,037 m/s. At Earth, the arrival V-infinity

is 5,416 m/s, resulting in an entry speed of 12,345 m/s,

the highest Earth entry speed for the launch period.

Total mission trip time is 573 days; and that varies only

1 day over the entire launch period. Total mission dV

for this 2035 mission is less than the total dV for the

2033 mission described previously.

9. Summary

With a trip time of 530 days, this 2033 crewed flyby

mission could serve as a precursor to a later Mars

surface mission. The approach outlined in this report

makes maximum use of existing launch systems (SLS,

Vulcan Heavy) and existing capsules (Orion, CST-100).

Two new developments would be required, the long

duration crew Habitat module and the Deep Space Burn

propulsive stage. The LO2/LH2 RL10 Engine TMI

stage discussed here could be a derivative of the

LO2/LH2 RL10 Engine SLS EUS stage already in

development. The 2033 crewed Mars Flyby mission can

be flown with three launches. Adding a fourth launch

provides two advantages: 1) a heavier Habitat sufficient

for 3 or 4 crew and increased levels of redundancy and

consumables, 2) the recovery of the Habitat back into

Earth orbit. Should the vehicle elements not be ready on

time for the 2033 opportunity, a suitable opportunity is

available two years later in 2035. The SLS, with its new

EUS upper stage, enables this mission and all

subsequent crewed Mars missions.

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10. Acknowledgements

The authors would like to thank Michael Elsperman,

Xavier Simon, Jerry Horsewood, David B. Smith,

Jonathan Median-Espitia, Andrew Chandler, Mike

Raftery, and Doug Cooke.

References

[1] Tobias, M.E., Griffin, D.R., McMillin, J.E., Haws,

T.D., and Fuller, M.E., “Booster Obsolescence and

Life Extension (BOLE) for Space Launch System

(SLS)”, IEEE Aerospace Conference 2020, Big

Sky, MT.

[2] Dunham, David, et al, “Earth-Moon Halo Orbit –

Gateway or Tollbooth?”, AAS 19-756, 2019.

[3] Z. R. Putnam, R. D. Braun, et al, “Entry System

Options for Human Return from the Moon and

Mars,” AIAA 5915, 2005

[4] Inspiration Mars, “Feasibility Analysis for a Manned

Mars Free-Return Mission in 2008”, Tito, Dennis;

MacCallum, Taber; Carrioc, John; May 8, 2013

[5] NASA Space Launch System (SLS) Mission

Planner’s Guide, ESD 30000 Revision A, Dec 2018

[6] Donahue, Benjamin; Boeing, “Crewed Lunar

Missions and Architectures Enabled by the NASA

Space Launch System” IEEE Aerospace 2020

Conference, Big Sky, Montana, March 2020

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Fig. 1 (top) SLS Evolution to Block-1B and Fig 2 (above) SLS Evolution to Block-2

Fig. 3 Launch Vehicle Fairing Size Comparison

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Figure 4 (top) 2033 Mars Flyby Trajectory Diagram, Fig. 5 (above) Timeline

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Table 1-A (Left)

2033 Mars Flyby Delta-Velocity Budget

(Crew Habitat Expended Option

Reference)

Table 1-B (Left)

2033 Mars Flyby Delta-Velocity Budget

(Crew Habitat Recovered Back at Earth

Option - Alternative)

Table 2 (above)

2033 Mars Flyby Earth Arrival Heating

Loads for Earth Entry Capsule

(Braun Ref 2)

Figure 6 (Left)

2033 Mars Flyby Earth Arrival Entry

Flight Corridor:

Altitude (km) vs Velocity (km/s)

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Fig. 8 (below)

Deployable Sun

Shield for Orion

Fig. 7 (Left)

2033 Mars Flyby

Habitat (right) and

Orion EEC (left)

Fig. 9 (left)

Crew EVA at

Mars Passage

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Figure 10 (Top) (Case 1)

Three-Launch Scenario: 3 SLS Bk-2, Off-loaded Orion to Mars, Crew Habitat: 18.3 mt

Figure 11 (Bottom) (Case 2) (DSB propulsion O2/CH4)

3-Launch: 2 SLS Bk-2, 1-Vulcan Heavy, Off-loaded CST100 to Mars: Crew Habitat: 15.7 mt

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Figure 12 (Case 3)

Three-Launch Scenario: 2 SLS Bk-2, 1-Vulcan Heavy, CST100 to EM-L2 only,

Small (High Entry Velocity) EEC (10 mt), Crew Habitat: 23.2 mt

Figure 13 (Case 4)

Four-Launch Scenario: 4 SLS Bk-2, Off-loaded Orion to Mars, Three NTO/MMH DSB Stages,

Crew Habitat: 28.1 mt

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Figure 14 (Case 5)

Four-Launch Scenario: 3 SLS Bk-2, 1-Vulcan Heavy, CST100 to EM-L2 only, two NTO/MMH

DSB Stages, Crew Habitat: 37.8 mt

Figure 15 (Case 6) (DSB Propulsion O2/CH4)

Four-Launch Scenario: 4 SLS Bk-2, Off-loaded Orion to EM-L2 Only, Three LO2/CH4 DSB

Stages, Crew Habitat: 26.4 mt Recovered Back to EM-L2

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Fig. 16 (above) and Fig. 17 (below) Mars Vehicle Approaching Mars

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Fig. 18 2035 Opportunity Mars Flyby Trajectory Diagram

Mars 2033 Human Flyby Mission

Abstract and Figures

Recommended publications

Existence and Multiple Solutions of the Minimum-Fuel Orbit Transfer Problem

The 2033 Crew Mars Flyby Mission

Design Concept For A Crewed Mars Flyby Mission in 2033

Capabilities for a 2033 Crew Mars Flyby Mission Launched with the NASA Space Launch System