The Space Launch System Capabilities with a New Large ...

The Space Launch System Capabilities with a New Large Upper StageBenjamin Donahue1 Sheldon Sigmon2

Boeing Exploration Launch Systems, Huntsville, AL 35824

The SLS is the most powerful rocket ever built and provides a critical heavy-lift launch capability enabling diverse deep space p p y p y g p p

missions. The exploration class vehicle launches larger payloads farther in our solar system, faster than ever before possible. A

new 8.4m Large Upper Stage (LUS), as a follow on to the interim Cryogenic Propulsion Stage (iCPS), can provide significant

increases in SLS payload injection capability. The new Large Upper Stage can be built at the Michoud Assembly Facility on the

same 8.4m tooling as the SLS Core stage. In this paper, SLS performance with both the iCPS and LUS are presented, and several

potential new missions are described.

Human Exploration

SLS enables exploration beyond LEO to scientifically valuable deep space destinations including our moon, asteroids and

ultimately Mars. It will launch the Orion Multi-Purpose Crew Vehicle (MPCV), granting unprecedented human access to new

space environments in the pursuit of knowledge and discovery. SLS boasts the most powerful propulsion system in the world,

significantly reducing crew travel time. The heavy-lift rocket is scheduled to take flight in 2017.

ScienceScience

SLS provides a critical new launch capability for planetary science and astronomy

missions expanding our understanding of the universe, the solar system and the

Earth. From large space telescope deployment to sample return missions, SLS’s

superior performance and 5 m to 10 m faring allows utilization of existing systems,

reducing development risks, size limitations and costs. The evolvable rocket

provides the greatest lift capacity system to dateprovides the greatest lift capacity system to date.

Security

SLS is a national security asset enabling the launch of large payloads for

intelligence, surveillance and reconnaissance missions. It provides a significant lift

capability directly to geostationary orbit augmenting current satellite deployment

options In addition to traditional military operations SLS can support globaloptions. In addition to traditional military operations, SLS can support global

asteroid capture or deflection missions to mitigate impact risks to known Earth

approaching objects.

Public-Private Partnerships

SLS is the first launch system in history capable of powdering humans, habitats

and space systems beyond our moon and into deep space Its flexible designand space systems beyond our moon and into deep space. Its flexible design,

superior performance and lift capability supports developing deep space markets

including space based solar power, extraterrestrial resource utilization and space

tourism beyond LEO. SLS reduces mission time and costs, supports a variety of

payloads and is scalable to address diverse exploration needs.

SLS/iCPS Block 1 Configuration

Copyright © 2013 by the Boeing Company. Published by the American Institute of Aeronautics and Astronautics, Inc., with permission.

Fig 1. SLS Block 1

SLS/iCPS Block 1 Configuration

Fig.1 - from top: Launch Abort System (LAS); Orion (MPCV), Orion spacecraft

adaptor (SA), Interim Cryogen Propulsion Stage (iCPS), Launch Vehicle

Spacecraft Adaptor (LVSA), Core stage, and two Five Segment Boosters (FSB).

1Boeing Defense, Space & Security (BDS), MC JV-08 Senior Member AIAA. 2BDS, MC JV-08, Advanced programs.

SLS Summary: Upper Stages and Performance

A new 8.4 meter diameter, Large Upper Stage (LUS) concept is presently under evaluation, data presented here is conceptual and

preliminary. The LUS can be built at the MAF on the same 8.4m tooling as the SLS Core stage. Because of increased thrust and

higher propellant loading increased payload injection capability can be achieved with the LUS A summary of missionhigher propellant loading, increased payload injection capability can be achieved with the LUS. A summary of mission

parameters for the SLS with the iCPS and with the LUS are listed in Table 1; column 1 data is for the initial Block 1 SLS/iCPS

configuration, and columns 2-4 contain data for three variants of the SLS/LUS configuration. BEO injection energies (C3) and

payloads for six destinations are also listed. The three versions of the LUS presented are identical except for the engines, thrust

vector control (TVC), feedlines and thrust structure. A variety of propellant loads were evaluated, but for this presentation LUS

usable propellant was set to 105mt (231,483 lbs). Tank size and propellant loads are identical for all three LUS variants. The LUS

at 264 000 lbs is about 3 5 times the mass of the smaller 5 0m diameter iCPS The three LUS versions are based on three engineat 264,000 lbs, is about 3.5 times the mass of the smaller 5.0m diameter iCPS. The three LUS versions are based on three engine

options, which include:

- the four RL10 C2 engine version (Col. 2), single engine thrust = 24,750 lbf, Isp vac = 462.5 s

- the dual MB-60 engine version (Col. 3), single thrust = 60,000 lbf, Isp vac = 465.0 s

-the single J2X engine version (Col. 4), single thrust vac = 294,000 lbf, Isp vac = 448.0 s

The MB-60 is a O2/H2 expander cycle concept engine that has never entered production.

Table 1. SLS Summary of Capabilities

Column 1 2 3 4

Name 1SLS / iCPS1xRL10B2

SLS / LUS4xRL10C2

SLS / LUS2xMB60

SLS / LUS1xJ2X

LEO final 2 100x975nmi 130x130nm 130x130nmi 130x130nmi

Booster x 2 3 FSB FSB FSB FSB

Core Stage 4 4xRS25 4xRS25 4xRS25 4xRS25

Upper Stage 5Interim Cryogenic

Propul Stage

LargeUpper Stg

LUS

LargeUpper Stage

LUS

Large Upper Stage

LUS

US diameter 6 5.0m 8.4m 8.4m 8.4m

US Usable Prop Load 7 27.1mt 105.0mt 105.0mt 105.0mt

Upper Stage PMF 8 0.883 0.900 0.900 0.895

Total Engine Thrust 9 24,750 lbf 99,000 lbf 120,000 lbf 294,000 lbf

Fairing diameter 10 5.0 m 8.4 m 8.4 m 8.4 m

Payload to LEO 11 70.0 mt 93.1 mt* 97.0* mt 105.2** mtPayload to LEO 11 70.0 mt 93.1 mt 97.0 mt 105.2 mt

Payload TLI C3= -2.0 12 24.0 mt 39.1 mt 39.7 mt 38.5 mt

Payload TMI C3= 11 13 20.2 mt 31.7 mt 32.6 mt 31.6 mt

Fairing dia Outer Planet 14 5.0 m 5.0 m 5.0 m 5.0 m

Payload Europa C3= 91 15 2.9 mt 8.1 mt 8.5 mt 7.1 mt


Payload Titan C3= 106 16 1.8 mt 5.7 mt 6.0 mt 4.6 mt

Payload Uranus C3= 137 17 0.13 mt 1.7 mt 2.0 mt 0.5 mt

*US stage off loaded; 63mt usable prop for LEO mission **95mt usable prop for LEO mission

Interim Cryogenic Propulsion Stage (iCPS)

SLS Upper Stages

In the following pages iCPS and LUS stages are described and payload performance, stage mass and other parameters are further

quantified. Data is preliminary. SLS mission payload performance listed here are for cargo missions.

The iCPS / Orion combination is pictured in Fig. 2, and the iCPS stage is pictured in Fig. 3. iCPS stage parameters are listed in

Table 2. The iCPS is a derivative of the Delta-IV 5m upper stage presently in production.

Table 2.

Interim Cryogenic Propulsion Stage (iCPS)

Figure 2. iCPS/Orion

Engines 1 x RL‐10 B2

Isp, vac 462.5 sec

Engine thrust, vac 24,750 lbf

Total thrust, vac 24,750 lbf

LH2 dia 5.0 m

LH2 propel 9,051 lbm

Intertank Composite X‐brace

LO2 dia 4.0 m

LO2 propel 52,949 lbm


RCS propel 700 lbm

Dry mass 8,445 lbm

Total mass 71,155 lbm

Fig. 3.

iCPS

Stage

SLS Block 1: iCPS Cargo Mission Performance

In Fig. 4 SLS/iCPS Block 1 vehicle injection energy (C3) is plotted vs payload. In Fig. 5 masses are listed for the SLS major

elements. For the Block 1 SLS/iCPS configuration, the Core stage delivers the iCPS and payload to a 20 x 975nmi orbit. From

there the iCPS fires to raise the perigee to 100nmi. From 100 x 975nmi the iCPS injects the payload to its destination. In Fig. 4

injection capability for the Delta IV Heavy EELV is also given for comparison

SLS iCPS

Delta-IV HeavyEuropa

Titan

injection capability for the Delta-IV Heavy EELV is also given for comparison.

TLI

Mars

Mars Free Return

LEO

Fig 4.

SLS/iCPS payload vs C3MISSION

Fairing 5.0m dia

Cargo Payload 52.9 klb

(24.0) mt

iCPS Total 71.2 klb

TLI C3 = -2

Usable prop 59.6 klb

Non -usable 2.4 klb

RCS prop 0.7 klb

Dry mass 8.4 klb

PMF 0.880

Interstage 8.4-5.0m

Fig. 5.

SLS / iCPS

Core Total 2,387.0 klb

Usable prop 2,116.9 klb

Non -usable 30.9 klb

Dry mass 220.6 klb

Booster FSB 1,607.8 klb

Startup prop 18.6 klb


major element

masses


Ejected inerts 9.0 klb

Slag 2.0 klb

Dry mass 211.3 klb

SLS with Large Upper Stage (LUS)

The SLS’s payload capability improves significantly by the addition of a new 8.4m Large Upper Stage (Fig. 6); payload increases

60% to TLI (39 vs 24mt) and 50% to LEO (105 vs 70mt) vs. the Block 1 vehicle. Additional payload enhances future science,

astronomy and human spaceflight missions. Fig. 7 list masses for the SLS/LUS. For the SLS/LUS mission, the Core stage stages

b bi ll d h i i i h i h bi h fi i i j h

Europa

TitanSLS LUS 1xJ2X

SLS iCPS 1xRL10

suborbitally, and the LUS ignites to continue the ascent to a 130 x 130 nmi LEO. From that orbit the LUS fires again to inject the

payload to its destination. In Fig 6 and 7 1xJ2X LUS data is shown. The LUS is described in more detail in the following pages.

TLI

Mars

Delta-IV Heavy

Mars Free Return

LEO

MISSION

Fairing 8.4m dia

Cargo Payload 85.8 klb

(38.5) mt

LUS Total 264.2 klb

TLI C3 = -2

Fig 6. SLS/LUS and SLS/iCPS

C3 vs. Payload

Usable prop 231.5 klb

Non -usable 3.7 klb

RCS prop 1.4 klb

Dry mass 27.6 klb

PMF 0.895

Interstage 8.4m

Fig.7.

SLS/LUS major

Core Total 2,388.4 klb


Non -usable 21.3 klb

Dry mass 222.0 klb

Booster FSB 1,607.8 klb

Startup prop 18.6 klb


j

element massesUsable prop 1,385.4 klb

Ejected inerts 9.0 klb

Slag 2.0 klb

Dry mass 211.3 klb

SLS Payload SummaryIn Fig. 8 payload is plotted vs C3 for LEO, TLI, Mars (TMI), Mars free return and Europa. The SLS provides a significant

increase in capability vs existing launchers. SLS/LUS can deliver 105.2mt to LEO, 38.5mt to TLI, 31.6mt to TMI, about

20.0mt to Mars free return and 7.1mt to Europa. Payload values are dependant on a multitude of factors all of which are not

LEO

Delta-IV Heavy

SLS / iCPS 1xRL10

SLS / LUS 1xJ2X

addressed in this report; values presented here are preliminary and subject to change.

TLI

Delta-IV Heavy

TLI

Mars

Europa

Mars Free Return

Upper Stage Gravity-loss vs Stack Thrust-to-Weight

LEO departure gravity (g)-loss is plotted vs stack T/W in Fig. 9. For the TLI mission, iCPS/Payload Thrust-to-weight (T/W) in

th 100 975 i LEO i 0 2 LUS/P l d T/W i th 130 130 i LEO i 1 2 f th 1 J2X i d 0 4 f th 4 RL10

Fig. 8. SLS Payload Summary

the 100x975nmi LEO is 0.2; LUS/Payload T/W in the 130x130nmi LEO is 1.2 for the 1xJ2X version, and 0.4 for the 4xRL10.


Figure 9. Upper Stage LEO departure g-loss vs stack T/W

SLS Large Upper Stage: Four RL10 Engine Configuration

The Large Upper Stage is an 8.4m diameter O2/H2 stage that can be built at the same MAF site, and on the same 8.4m diameter

tooling, as the SLS Core stage. Building the LUS and Core together offers several economic advantages; principally commonality;

common elements (tank domes, avionics as examples), personnel, processes, testing, and qualification. Three LUS options are

illustrated in Figs. 8-11. The first LUS option presented, the 4xRL10 version, is shown in Fig. 8. All the LUS variants feature a

8.4m LH2 tank and a 5.5m LO2 tank. A range of LUS propellant loads were evaluated; for this report a usable load of 105mt

(231,488 lb) was selected. This 4xRL10 LUS weighs 262,752 lb fully fueled. The Core interstage (8.4m) attaches to the LUS at

the base of the 8.4m LH2 tank aft skirt. The 4 RL10 C1 engines together produce 99,000 lbf thrust.

blTable 3.

Large Upper Stage‐ 4 x RL10 C2

Engines 4 x RL‐10 C2

Isp, vac 462.5 sec

Thrust, vac 24,750 lbf


LH2 dia 8.4 m

LH2 propel 34,334 lbm

Intertank Composite X‐brace

LO2 dia 5.5 m


RCS propel 1,432 lbm

Dry mass 26,133 lb

Total mass 262,752 lb

Fig. 8.

4xRL10

Large

UpperUpper

Stage

concept


Figure 9. Exploded view

SLS Large Upper Stage: 4 x RL10 Configuration

Fig 10 (left)

LUS: Propellant Load

Payload is plotted vs LUS usable

propellant in Fig. 10 for the

4xRL10 LUS. TLI Payload is

maximized at 105mt load

(bottom curve). When flown to

LEO rather than TLI, off-loading

to 63mt prop maximizes LEO

payload (top curve). Due to its

higher thrust, LEO payload

maximizes at a higher prop load

for 1xJ2X version LUS.

Fig 11a.

4xRL10

LUS in

Orbit

Figure 11b.

Large Upper


Large Upper

Stage concept

Elements

SLS Large Upper Stage: Single J2X Engine Configuration

The single J2X version of the LUS is illustrated in Fig. 12. This LUS also

utilizes a 8.4m dia LH2 tank and a 5.5m dia LO2 tank (Table 4). This

Table 4.

Large Upper Stage‐ 1 x J2Xutilizes a 8.4m dia LH2 tank and a 5.5m dia LO2 tank (Table 4). This

stage is identical to the 4xRL10 version with the exception of the engines,

TVC, feedlines and aft thrust structure. This LUS weighs 264,211 lbm

fully fueled. The thrust of 1xJ2X is 3 times the thrust of 4xRL10s. The

1xJ2X LUS’s thrust advantage allows it to lift more to LEO than the

4xRL10 version. However, due to its lower Isp (448 vs 462 sec), it

delivers less payload to high injection C3 destinations like Europa (7.1 vs

Engines 1 x J2X

Isp, vac 448.0 sec


LH2 / LO2 dia 8.4 m / 5.5m

LH2 propel 34,334 lbmdelivers less payload to high injection C3 destinations like Europa (7.1 vs

8.1mt). The J2X engine has a mass of 5,450 lb, by comparison, the RL10

C1 engine weighs 664 lb each, or 2,656 lb for four.


RCS propel 1,432 lbm

Dry mass 27,593 lb

Total mass 264,211 lb


Figure 12. 1xJ2X Large Upper Stage concept

SLS Missions:

SLS will play a critical part in enabling the next steps in human space exploration beyond Earth orbit. In the following pages several

representative missions are presented, each launched with the SLS.

Mission to Translunar Space

The SLS will enable the next steps in human space exploration beyond Earth orbit to translunar space. Building a translunar outpost

is an important first step in retrieving an asteroid, returning to the moon or venturing to Mars. NASA’s phased development plan to

evolve lifting capability will allow larger payloads to be launched economically, opening new options for larger vehicles. The SLS

will be ideally suited to deliver a variety of payloads to translunar space. A SLS/LUS could launch multiple translunar elements

simultaneously, reducing the number of overall launches and decreasing the time required to assemble a translunar outpost.

Translunar missions will be undertaken once the SLS and Orion are available

as launch and crew vehicles. Crew operations in translunar space could be

significantly enhanced by providing additional systems and EVA capabilities

beyond those available from Orion only missions. An Exploration Platform

located in translunar space would improve the science and technical return of

the early missions while also increasing Orion capability through resource

provision and providing an abort location and safe haven for vehicle

contingencies. Overall this would increase mission duration, expand mission

utilization option and enhance safety.

Ideally, to increase international participation and share costs, an Exploration

Platform would be created using existing hardware from a number of sources.

Russian systems are well developed and ideal for these new uses, such as

adapting current Russian Science Power Module (SPM) and node designs for

translunar use. Hardware from the Space Shuttle and International Space

Station (ISS) programs, such as the Orbiter Docking System (ODS) and the ISS

Multi-Purpose Logistics Module (MPLM), could be combined with existing

satellite hardware.

Fig. 14 shows the launch of a Russian SPM-derived core module (lower) with

an American ODS-based utility module (upper). Boeing has coordinated with

RSC Energia to study how these elements might be applied in translunar space

for asteroid operations. RSC Energia concluded that a new hybrid module

might be the best option. This module would shorten the pressurized section of

the SPM and add a new node/docking ball section. The utility module is built

around an existing ODS by adding flight-proven Boeing 702 satellite systems

to create an independent, fully functional vehicle. The utility nodule provides

two NASA Docking Standard systems for docking operations with visiting

vehicles.

These two elements together would provide all the key functions required for a

crew base vehicle: EVA capability, internal crew volume, capability to receive


cargo missions, Orion mission extension and future extensibility. Either of the

elements is capable of sustaining the overall vehicle with power and attitude

control. The SPM provides the bulk of the crew habitable volume in the base

vehicle and the utility module provides the important EVA capability.

Fig. 14. SLS/LUS Launch with Russian

SPM and US Utility Module

SLS Mission: Exploration Platform

Once the SPM-derived core module and utility module are in place, a second SLS / LUS launch would deliver an Orion with crew

and a habitat module, as shown in Fig.15. Following launch and translunar insertion, the Orion would separate, turn 180 degrees and

then dock to the habitat module The upper stage would separate and the Orion would fly the habitat module to the SPM-derivedthen dock to the habitat module. The upper stage would separate and the Orion would fly the habitat module to the SPM derived

core module /utility module already in operation. The habitat module is based on an existing MPLM remaining from the ISS

program. The module would be refurbished to take better advantage of the internal volume than was possible with the ISS

removable rack concept. The berthing mechanism would be replaced with an NDS port and an additional NDS port would be added

at the opposite end. Internally, the habitat would be outfitted with crew living quarters, medical services, exercise equipment, a

galley and payload facilities, as well as additional core systems hardware for redundancy.

As a whole, these three elements—SPM-derived core module, utility

module and habitat—form a very capable Exploration Platform that

would enable and significantly enhance human operation in translunar

space. Initial missions would focus on science, perhaps by studying a

retrieved asteroid. Ongoing research would focus on the deep space

environment, particular the radiation environment and its effect on

humans. Later, the Exploration Platform would serve a base or

jumping-off point for missions to the moon or Mars. The Exploration

Platform would be ideally located to support lunar surface operations,

such as through the remote operation of robotic exploration or

construction vehicles. The Exploration Platform could also be used as

an assembly location for large Mars vehicles and a transfer point for

crew on their way to Mars.

Using the SLS/LUS would allow the Exploration Platform to be

constructed and crewed in only two launches, as opposed to the four

missions required using SLS/iCPS, thus saving cost and significantly

shortening the time required to start accruing the benefits of a crewed

Exploration Platform in translunar space. The SLS family opens up

new possibilities for human space exploration beyond Earth, including

missions to translunar space (Fig. 16) and Mars.


Fig. 15. SLS/LUS with Orion and Hab Module Fig. 16. Exploration Platform enables future missions

SLS Mission: Large Crew Habitat EmplacementThe Bigelow BA-2100 is a stand-alone, self-sufficient module for long duration human habitation. It has all the services and

resources (propulsion, power generation, attitude control, etc.) to support a wide variety of missions and purposes. In LEO

for commercial operation, it could serve as an orbital hotel. It could also be the base module for additional expansion, for

instance to add scientific research modules. The BA-2100 is an inflatable module that is deployed after launch to provide

2100 cubic meters of pressurized volume. The module is shown in uninflated in Fig. 17 and 18.

SLS allows delivery of the BA-2100 via direct insertion to a low earth orbit

and is the only launch vehicle capable of delivering a payload this large to

Fig. 17. Habitat and Upper Stage in LEO

and is the only launch vehicle capable of delivering a payload this large to

LEO. SLS provides significant mass margin that can be used for additional

crew consumables or water for radiation protection or additional payloads

The BA-2100 is shown packaged in a 10 m fairing for launch (Fig. 18). After

launch and separation from the first stage, the upper stage would move the

BA-2100 into the desired orbit and separate as well The BA-2100 would thenBA 2100 into the desired orbit and separate as well. The BA 2100 would then

be fully activated to deploy the solar power and thermal arrays and inflate (Fig.

25). Once fully operational, crew visits could commence, such as with the

Boeing CST-100 as shown in Fig. 19.


Fig. 18. SLS Habitat launch configurationFig. 19. Habitat at destination

SLS Mission: Uranus Concept

SLS / Large Upper Stage Injection Mass to Uranus (C3=131 km2/2) is 1.7mt.

Mission Objective -

-Deliver a small payload into orbit around Uranus and a shallow probe into the planet’s atmosphere

Mission Rationale -

-Investigate the ice giant system’s atmospheric and magnetic properties, determine the distribution of thermal

emission from the planet’s atmosphere, refine the gravitational harmonics of the planet and conduct close flybys of

any large satellites. Representations of the spacecraft are given in Fig. 20 and 21.

SLS Capabilities -

-SLS mission design shortens travel time, allows for the inclusion of increased mass and removes the need for a

solar electric propulsion stage (SEP) simplifying overall spacecraft and mission design.

Fig. 20. Uranus Mission Spacecraft


Fig. 21. Uranus Mission Spacecraft

SLS Mission: Solar Probe 2 Concept

Mission Objective -

-Launch the first spacecraft capable of frequent and close encounters with the sun

Mission Rationale -

-Solar Probe 2 will provide researchers both in-situ measurements and imagery supporting corona heating and solar

wind acceleration investigations. It will also be part of the spacecraft fleet charged to develop the critical

forecasting capability of the space radiation environment in support of human and robotic exploration.

SLS Capabilities -

-SLS mission design incorporates the advantages of both the Solar Probe and Solar Probe Plus spacecraft. It

provides a low perihelion distance (as low as 5 solar radii) and frequent revisit times without the use of

radioisotope thermoelectric generators. Illustrations of the spacecraft are given in Fig. 22 and 23.

Fig. 22. Solar Probe 2 Spacecraft in Route


Fig 23. Solar Probe 2 Spacecraft

SLS ATLAST Space Telescope Concept

Mission Objective -

-Characterize Exoplanets and search for signs of life

Best option for extrasolar life-finding facility

Observe ~85 stars 3 times each in a 5-year period

–Probe super massive black holes (SMBH)

Direct measurements of the mass of high redshift SMBH

–Exploration of the Modern Universe

Enable star formation histories to be reconstructed for hundreds of galaxies

Track how and when galaxies assemble their present stars

–Constrain dark matter

Measure the mean density profile of dwarf spheroidal galaxies (dSph), a fundamental constraint on the

nature of dark matter

Mission Rationale -

SLS offers the possibility in a single launch of an 8m Monolithic or a 16m deployable ATLAST

–Telescope deployed at Earth – Sun L2

–Human and/or robotic servicing would be highly desirable extending the life up to 20-30 years

–10 times the resolution of JWST and up to 300 times the sensitivity of the HST

–A monolithic aperture is better than a segmented aperture

-JWST is using a segmented, deployed mirror architecture only because it is the only way to launch a 6.5

meter aperture observatory with a 4.5 meter diameter rocket

-A monolithic mirror can achieve diffraction limited performance at a shorter wavelength than a segmented

mirror with much difficulty, complexity, cost and risk.

SLS Capabilities -

Without the SLS multiple EELV launches and in space assembly are required for the 16m version and no other

launch vehicle is capable of launching an 8m Monolithic telescope


Fig. 24 ATLAST space Telescope concept

Summary

The SLS provides a critical heavy-lift launch capability enabling diverse deep space missions. The exploration class p y p y g p p p

vehicle launches larger payloads farther in our solar system, faster than ever before possible. This added payload to

destination that can be provided by a new Large Upper Stage would be an enhancement for future science, astronomy

and Human spaceflight missions. The Large Upper Stage can be built at the Michoud Assembly Facility on the same

8.4m tooling as the SLS Core stage and achieve the economic benefits that come with commonality of subsystems,

processes and personnel. The SLS in its evolving configurations will enable a broad range of exploration missions.

SLS SLS

iCPS LUS J2X

Payload mt Payload mt Increase

LEO 70.0 105.2 50 %

Lunar TLI 24.0 38.5 60 %

Mars TMI 20.2 31.6 56 %

Europa 2.9 7.1 144%


The Space Launch System Capabilities with a New Large ...

Documents

Transcript of The Space Launch System Capabilities with a New Large ...