JuanCastillo29/space-mission-intelligence-data huggingface dataset:
The SLS employs a Right-Handed Cartesian Body-Fixed coordinate system, with a heritage orientation derived from the Space Shuttle Program, as shown in Figure 3-3.
 +XSLS points aft down the vehicle axis.  +YSLS points to the centerline of the right-hand (RH) booster.  +ZSLS completes the right-hand rule.

[Figure: Figure 3-3. SLS Vehicle Coordinate System]

Note: RH booster is on the right side of the stack (the +YSLS side of the launch vehicle) when looking at SLS with the Mobile Launcher Tower beyond. In rendered views of the vehicle, the – YSLS (Left) booster features a black band below the nose cone.
Orion or existing 5-m class diameter Payload Fairing (PLF) for early cislunar and deep space
applications. The initial SLS Block 1 test flight, EM-1, is currently planned to accommodate an
uncrewed Orion and a number of SPL. Its purpose is to test SLS launch capabilities and Orion’s
ability for safe translunar crew return; this is planned to be available for launch no earlier than 2020.
An overview of the SLS Block 1 crew configuration is shown in Figure 3-4. The cargo version of
the Block 1 vehicle uses a 16.7 ft (5.1 m) diameter PLF planned for Science Mission-1 (SM-1)
which will accommodate an interplanetary Europa Clipper launch no earlier than 2022. The cargo
configuration shown in Figure 3-5 does not provide SPL accommodations and will not be available
after the introduction of SLS Block 1B vehicle; it is provided here for reference only.
Michoud Assembly Facility, where the Saturn stages and shuttle external tanks were manufactured.
This 27.6 ft (8.4 m) diameter, 200 ft (61 m) long tank forms the rocket’s structural backbone. The
the vehicle-level avionics. Four RS-25 engines will power the SLS Core Stage. These human-rated
engines support the SLS goal of safety, with a record of 100 percent mission success for the engines
over 135 flights and has accumulated over 1 million seconds of ground hot-fire experience.
pair of SRBs, also of Space Shuttle Program heritage. The SLS Program is leveraging research,
development and testing conducted under the Ares Program to upgrade the SRBs from the four-
segment version flown on the shuttle to more powerful five-segment boosters. Shuttle-heritage
hardware and design includes forward structures, metal cases, aft skirts and thrust vector control
elements. The upgraded hardware and expendable design include solid rocket motors, avionics and
asbestos-free insulation.
The electronic version is the official approved document.

SLS Block 1 Integrated Spacecraft/Payload Element (ISPE). The Block 1 crew configuration ISPE is composed of the Launch Vehicle Stage Adapter (LVSA), ICPS, and SLS Orion Stage Adapter (OSA) as shown in Figure 3-4 and 3-6. The OSA provides SPL accommodations as needed during Orion missions. For the Block 1 cargo configuration ISPE, the OSA is replaced with a Commercial Off the Shelf (COTS) 5 m-class diameter payload adapter and PLF as shown in Figure 3-5. The payload adapter accommodates PPL as needed. In-space propulsion for SLS Block 1 payloads is provided by the ICPS, derived from an existing United Launch Alliance (ULA) Delta Cryogenic Second Stage (DCSS). This 16.7 ft (5.1 m) diameter stage has flown on more than 20 launches of the Delta IV Evolved Expendable Launch Vehicle (EELV). A single RL10 LOX/LH2 engine powers the stage which has a mission life of approximately 8 hours. It interfaces to the SLS Core Stage via the LVSA. The ICPS 16.7 ft (5.1 m) diameter forward skirt provides a standard interface to the Orion via the SLS OSA for crewed missions and interfaces to available 5 m-class diameter payload adapters/PLFs for a cargo mission.
ICPS Launch Vehicle Stage Adapter (LVSA)
Orion Stage Adapter
18.0’ (5.5 m)
16.7’ (5.1 m)
While the SLS Program prepares for the first SLS Block 1 flight (EM-1), development work has already begun for evolution to the higher-performance SLS Block 1B configuration. The development of the new EUS will be required in order to ultimately provide planned TLI capabilities. For Block 1B, the configuration of the Core Stage will be used with only minor subsystem changes from Block 1. This commonality-based strategy will reduce the cost and risk of Block 1B development by maintaining similar interfaces to flight hardware and ground systems as those used for Block 1. Figure 3-7 details the Block 1B crew configuration; it is planned to be first available to support EM-3, a crewed Orion flight in support of lunar Gateway assembly no earlier than 2024. Figure 3-8 details the Block 1B cargo configuration using a 27.6 ft (8.4 m) diameter PLF.
Verify this is the correct version before use.

SLS Block 1B Core Stage. Core Stage will remain essentially unchanged from the Block 1 configuration.
SLS Block 1B Solid Rocket Boosters. The Boosters will remain essentially unchanged from the Block 1 configuration.
SLS Block 1B Exploration Upper Stage (EUS). The Block 1B EUS shown in Figure 3-9 is being developed to provide both ascent/circularization and in-space transportation for payloads. Four RL10 LOX/LH2 engines power the stage. Onboard batteries provide electrical power, while a passive thermal control system minimizes cryogenic propellant boil-off during the EUS lifetime. Mechanical interface to the SLS Core Stage is through a 27.6 ft (8.4 m) diameter Interstage. The 27.6 ft (8.4 m) diameter EUS Forward Adapter interfaces to the 8.4m-class diameter Payload Adapter (PLA).
EUS Interstage USA / PLA
27.6’ (8.4m) 18.0’ (5.5m)
[Figure: Figure 3-9. SLS Exploration Upper Stage]

SLS Block 1B ISPE. The Block 1B crew configuration ISPE is composed of a SLS PLA and USA (Figure 3-7) with CPL and SPL being accommodated on the PLA as needed. The USA then provides a standard interface to the Orion Spacecraft Adapter (SA). The Block 1B cargo configuration ISPE accommodates a PPL (and SPL as needed) using a SLS PLA and PLF (Figure 3-8).
[Figure: Figure 3-11 illustrates the progress to date made to support the initial launch (EM-1) and test of Block 1 SLS and its associated ground infrastructure. The SLS OSA shown in the upper right depicts the installation of SPL dispensers to support delivery to a cislunar destination and beyond. Three Block 1 flights in the early 2020’s are currently planned to demonstrate critical capabilities to be leveraged for future SLS Block 1B flights beginning in the mid-2020s.]

[Figure: Figure 3-11. SLS Block 1 Progress in Support of First Flight (EM-1)]
[Figure: Figure 4-1. NASA’s Exploration Campaign Overview]

The goal of NASA’s Exploration Campaign is to lead an innovative and sustainable program of exploration with commercial and international partners to enable human expansion across the solar system and to bring back to Earth new knowledge and opportunities. Beginning with missions beyond LEO, the United States will lead the return of humans to the Moon for long-term exploration and utilization, followed by human missions to Mars and other destinations. This campaign has the following steps:
• Transitioning LEO into commercial activities
– International Space Station operations, crew and cargo delivery • Developing a cislunar Gateway and planning for lunar surface landings
– Focus on the lunar surface for science and exploration
– Advancing planetary lander capabilities
– Honing sample return operations – lunar and Mars samples
– Deploying CubeSats, small sats and other science platforms
– Provides opportunities for commercial ventures and international partnerships
– Informing Mars transportation, life support systems, human health and research • Leveraging industry spaceflight capabilities and international partnerships for long-term solar system exploration
Jettison�@�3-sigma�<0.1� BTU/ft2-sec

| Representative |  | Lunar Injection (C3= -2.0 km2/sec2) |  |
| --- | --- | --- | --- |
| Event |  | Altitude | Velocity |
| EUS TLI MECO | 6,900 nmi (12,800 km) |  | 3.9 MPS (6.3 km/sec) |
| EUS TLI + 1hr | 15,400 nmi (28,600 km) |  | 2.9 MPS (4.6 km/sec) |
| EUS TLI + 8 hr | 32,800 nmi (60,800 km) |  | 1.9 MPS (3.1 km/sec) |
adjustments for targeting, may affect performance and ground impact zones. While fairing
separation varies depending on payload requirements, vehicle performance and atmospheric
heating during ascent, the fairing is typically jettisoned during the Core Stage burn when the free
depressed trajectory, the fairing may be carried beyond the staging event and jettisoned during the
upper stage burn. The separation of the SRBs, and then later Core Stage separation, will occur at
different trajectory locations during ascent due to the rate of fuel consumption and differences in
payload mass and target orbits.
The electronic version is the official approved document.

For SLS Block 1B/2, the EUS performs the final portion of the ascent burn and injects itself and the associated payload into some type of Earth orbit. The EUS can loiter here for a determined amount of time before performing its final burns; loiter time is dependent upon the mission's propulsive needs balanced against the degree of propellant boil-off incurred in LEO. The final EUS burn could be to raise the payload to a higher Earth orbit. For SLS crew mission CPL and SPL operations begin after Orion has separated from the EUS/USA. Depending on mission requirements, this can occur approximately five to eight hours from launch.
SLS cargo mission PPL can separate from the upper stage post-PLF separation. Depending on mission requirements, this can occur approximately one to eight hours from launch.
For LEO staging scenarios, the EUS could fly to an elliptical orbit and then if needed perform a circularization/perigee burn to establish the payload in LEO. The EUS would then re-enter Earth's atmosphere post-payload separation for disposal. Another alternative is that the EUS could insert the payload directly into a circular orbit followed by EUS disposal.
Lunar Vicinity

For typical lunar trajectories shown in Figure 4.2, the SLS upper stages loiter in LEO for a predetermined amount of time before performing its TLI burn. The amount of time in orbit depends on the mission's propulsive needs balanced against the degree of propellant boil-off incurred in the orbit. The upper stage then performs an injection burn to initiate a lunar transfer orbit. For SLS crew mission CPL and SPL operations begin after Orion has separated from the EUS/USA. Depending on mission requirements, this can occur approximately five to eight hours from launch.
SLS cargo mission PPL can separate from the upper stage post-PLF separation. Depending on mission requirements, this can occur approximately one to eight hours from launch.
For the uncrewed SLS test flight (planned for EM-1), the SLS Block 1 ICPS will perform a TLI maneuver to enable Orion to insert into a Distant Retrograde Orbit (DRO) around the Moon, as shown in Figure 4-4. After Orion separation, the ICPS continues on a heliocentric disposal trajectory. During this period, the SPLs can be deployed at various locations along the trajectory, as shown in Figure 4-5. The “bus stops” in Figure 4-5 are representative release points only; deployment at other locations along the disposal path is possible.
The electronic version is the official approved document.

[Figure: Figure 4-4. SLS Block 1 Distant Retrograde Orbit (DRO) Trajectory (EM-1)]

[Figure: Figure 4-5. Representative Secondary Payload Jettison “Bus Stops”]

For the first SLS crewed mission (planned for EM-2), the SLS Block 1 ICPS will insert Orion into an elliptical orbit to support final checkout of the spacecraft, After this, Orion flies a Hybrid Free
The electronic version is the official approved document.

Return trajectory around the Moon, as shown in Figure 4-6. Following separation from Orion, ICPS will put itself on a heliocentric disposal trajectory. Deployment of any SPLs at different ICPS locations may occur during this period, as shown in Figure 4-5.

[Figure: Figure 4-6. SLS Block 1 Hybrid Free Return Trajectory (EM-2)]

For typical SLS Block 1B crew missions to cislunar destinations (e.g., Lunar Gateway), the SLS EUS will perform a TLI burn to enable Orion to insert into a Near Rectilinear Halo Orbit (NRHO), shown in Figure 4-7. Following separation from Orion, EUS will continue onto a heliocentric disposal trajectory. Deployment of any SPLs may occur during this period as shown in Figure 4-5.
stage injects itself and the associated payload into a circular or elliptical orbit. The upper stage can
loiter in this orbit for a predetermined amount of time before performing its final injection burn.
Time in orbit is dependent upon propellant load, balanced against the degree of propellant boil-off
incurred in the orbit. The SLS upper stage then performs an injection burn to initiate escape from
Earth’s gravitational influence. Depending on mission requirements, this can occur any time
between one and eight hours after launch.
SLS Mass Delivery Performance Definitions

Due to the range of potential SLS payload accommodations, it is important for the user to understand SLS accounting of performance to destination when determining “useful” performance available to spacecraft/payload. For the purposes of this document, the following SLS mass delivery definitions should be used for cargo missions, as shown in Figure 4-8.
 Injected Mass at LEO (IMLEO) o Includes upper stage dry mass, unused upper stage fuel on-orbit and Payload System Mass (PSM)  Payload System Mass o Cargo configuration payload capability o PSM = IMLEO - Upper Stage Burnout Mass o Includes mass of spacecraft/payload and associated PLAs required to interface to upper stage  Useful PSM* o Useful PSM = PSM – Program Manager’s Reserve (PMR) o (See Section 4.2.2 for SLS PMR approach)  Spacecraft/payload – user-provided item delivered by SLS to inspace destination
*Referred to as Useful Load Mass (Delta IV) & Payload System Weight (Atlas V)
Performance relative to spacecraft/payload delivery in conjunction with crew missions (e.g., CPL) is handled differently than ELV flights due to the addition of the USA. See Section 4.2.5 for CPL performance definitions.
SLS Performance Margin and Reserve Approach

Since SLS Block 1B and Block 2 configurations are still in development, all performance estimates include appropriate SLS vehicle performance reserves. Payload performance identified in this document already takes these margins into account; users should not subtract these margins from published performance metrics.
Mass Growth Allowance (MGA) accounts for the inability to accurately predict hardware mass before it is fully designed and constructed. This margin is added to hardware mass estimates. The amount of MGA carried varies from system to system, according to the level of fidelity of each area of the design.
Program Manager’s Reserve (PMR) is a performance allocation that is held in reserve to address unexpected events during development, such as a change to the vehicle requirements. The MGA
The electronic version is the official approved document.

and PMR will be maintained throughout the development of SLS to ensure performance will meet program goals despite potential increases in SLS inert mass or if the performance of its propulsion systems is lower than expected.
Flight Performance Reserve (FPR), in addition to these development phase margins, is carried as an operational margin. This performance allocation is held in reserve by vehicle management to account for day-to-day variation in launch environments and the allowable variation in hardware performance.
SLS Earth Orbit Performance

SLS capabilities are optimized for missions to deep space rather than LEO. While there are currently no plans to use SLS missions to LEO, Figures 4-9 through 4-12 provide LEO capabilities as a reference. SLS Useful Payload System Mass to various Earth orbits is given in Figures 4-9 and 4-10 for circular orbits and Figures 4-11 and 4-12 for elliptical orbits. Unless otherwise noted, all orbits are to a 28.5 degree inclination.
Performance for the SLS Block 1 to elliptical orbits is based on using a 16.4 ft (5.1 m) diameter PLF that is 62.7 ft long (19.1 m) as a maximum volume case. Use of fairings shorter than this, for example a 47 ft (14.3 m) long PLF, would increase performance available to payloads.
Performance for the SLS Block 1B is based on using a 27.6 ft (8.4 m) diameter PLF that is 62.7 ft long (19.1 m); such a configuration is also known as the SLS “8.4m PLF, Short”. Both minimum and maximum performance cases are shown for Block 1B, representing a projected performance range based on current configuration development studies.
Performance for the SLS Block 2 is based on using a 27.6 ft (8.4 m) diameter PLF that is 90 ft long (27.4 m); such a configuration is also known as the SLS “8.4m PLF, Long”. It represents a reasonable minimum performance available to payload at this time. Use of fairings shorter than the 8.4m PLF, Long on SLS Block 2 would increase performance available to payloads.
Since Core Stage disposal locations are still being assessed, final optimization will include Core Stage disposal targeting, which may affect performance predictions. More specific performance capability can be evaluated on a case by case basis; contact the SLS SPIE office at nasaslspayloads@mail.nasa.gov for more details.
(lbm)
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(lbm)
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(lbm)
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220 199985 214222 230799 1500 148898 160641 174455 10000 73516 80233 87559 300 194461 209417 225991 2000 136393 147285 160406 15000 64538 70043 75799 400 189424 204025 220308 2500 126319 136545 149132 20000 59832 64532 69667 500 184732 198947 214956 3000 118074 127757 139916 30000 55028 58805 63361 600 180202 194156 209905 4000 105442 114303 125801 40000 52217 55567 59809 800 171960 185336 200599 5000 96284 104558 115457 50000 49980 53038 57060 1000 164545 177402 192213 7500 81796 89155 98172

[Figure: Figure 4-9. SLS Block 1B/2 Useful PSM to Circular Earth Orbits (English)]
(t)
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(t)
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(t)

(km)
(t)
(km)
(km)
(t)
(t)
(t)
407 90.7 97.2 104.7 2778 67.5 72.9 79.1 18520 33.3 36.4 39.7 556 88.2 95.0 102.5 3704 61.9 66.8 72.8 27780 29.3 31.8 34.4 741 85.9 92.5 99.9 4630 57.3 61.9 67.6 37040 27.1 29.3 31.6 926 83.8 90.2 97.5 5556 53.6 57.9 63.5 55560 25.0 26.7 28.7 1111 81.7 88.1 95.2 7408 47.8 51.8 57.1 74080 23.7 25.2 27.1 1482 78.0 84.1 91.0 9260 43.7 47.4 52.4 92600 22.7 24.1 25.9 1852 74.6 80.5 87.2 13890 37.1 40.4 44.5

[Figure: Figure 4-10. SLS Block 1B/2 Useful PSM to Circular Earth Orbits (SI)]
(nmi)
(lbm)
(nmi)
(lbm)
(nmi)
(lbm)
220 201001 2000 163776 15000 88433 300 199046 2500 156132 20000 81394 400 196693 3000 149278 30000 73815 600 192252 4000 137327 40000 69819 800 188081 5000 127832 50000 67360 1000 183854 7500 111316 75000 64004 1500 173601 10000 100877 100000 62293
(lbm)
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(lbm)
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220 199284 214188 231262 2000 164463 176204 191107 15000 107046 115081 127210 300 197828 211843 228785 2500 157980 169290 183822 20000 101557 109244 121189 400 195157 208987 225768 3000 152429 163373 177596 30000 95425 102728 114476 600 190156 203641 220085 4000 143420 153774 167516 40000 92084 99177 110745 800 185568 198737 214906 5000 136422 146323 159709 50000 89979 96943 108352 1000 181345 194225 210140 7500 124273 133395 146201 75000 87049 93835 104965 1500 172133 184389 199746 10000 116473 125102 137567 100000 85528 92222 103183

[Figure: Figure 4-11. SLS Block 1/1B/2 Useful PSM to Elliptical Earth Orbits (English)]
(km)
(t)
(km)
(t)
(km)
(t)
80.0
407 91.2 3704 74.3 27780 40.1 556 90.3 4630 70.8 37040 36.9 741 89.2 5556 67.7 55560 33.5 1111 87.2 7408 62.3 74080 31.7 1482 85.3 9260 58.0 92600 30.6 1852 83.4 13890 50.5 138900 29.0 2778 78.7 18520 45.8 185200 28.3
60.0
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(km)
(km)
(km)
(t)
(t)
(t)
407 90.4 97.2 104.9 3704 74.6 79.9 86.7 27780 48.6 52.2 57.7 556 89.7 96.1 103.8 4630 71.7 76.8 83.4 37040 46.1 49.6 55.0 741 88.5 94.8 102.4 5556 69.1 74.1 80.6 55560 43.3 46.6 51.9 1111 86.3 92.4 99.8 7408 65.1 69.8 76.0 74080 41.8 45.0 50.2 1482 84.2 90.1 97.5 9260 61.9 66.4 72.4 92600 40.8 44.0 49.1 1852 82.3 88.1 95.3 13890 56.4 60.5 66.3 138900 39.5 42.6 47.6 2778 78.1 83.6 90.6 18520 52.8 56.7 62.4 185200 38.8 41.8 46.8

[Figure: Figure 4-12. SLS Block 1/1B/2 Useful PSM to Elliptical Earth Orbits (SI)]
The electronic version is the official approved document.

[Table: Table 4-1. Useful PSM to Earth Escape]

| C3 Actual (km2/s2) | Block 1 5 m x 19.1 m Fairing (lbm) (t) | Block 1B 8.4 m x 19.1 m Fairing Min. Payload Max. Payload (lbm) (t) (lbm) (t) |  | Block 2 8.4 m x 27.4 m Fairing (lbm) (t) |
| --- | --- | --- | --- | --- |
| -20 -10 -0.99 0 10 20 30 40 50 60 70 80 83 90 100 110 120 130 140 | 59911 27.2 58663 26.6 48045 21.8 39571 17.9 32708 14.8 27100 12.3 22440 10.2 18537 8.4 15234 6.9 12415 5.6 11650 5.3 9989 4.5 7885 3.6 6051 2.7 4440 2.0 3017 1.4 1755 0.8 | 111503 50.6 95463 43.3 82848 37.6 81614 37.0 69570 31.6 59036 26.8 49666 22.5 41270 18.7 33782 15.3 27117 12.3 21183 9.6 15896 7.2 14424 6.5 11177 5.1 6958 3.2 3175 1.4 | 120671 54.7 103470 46.9 89988 40.8 88609 40.2 75681 34.3 64204 29.1 53908 24.5 44742 20.3 36612 16.6 29408 13.3 23024 10.4 17358 7.9 15784 7.2 12319 5.6 7826 3.5 3811 1.7 213 0.1 | 132241 60.0 113949 51.7 99151 45.0 97589 44.3 82787 37.6 69663 31.6 58163 26.4 48133 21.8 39401 17.9 31793 14.4 25147 11.4 19319 8.8 17708 8.0 14188 6.4 9653 4.4 5629 2.6 2044 0.9 |

Co-manifested Payload Missions. SLS Block 1B/2 can simultaneously transport Orion and CPLs using a PLA/USA to lunar destinations as needed. While SLS Block 1B performance capability for CPL is still under development, the current CPL TLI delivery mass goal for the first Block 1B flight (planned for EM-3) is 17,637 lb (8.0 t) based on an overall delivery of 80,028 lb (36.3 t) to a cislunar destination. The current CPL TLI delivery mass goal for the second and all subsequent Block 1B flights is 19,842 lb (9.0 t) based on an overall delivery of 82,232 lb (37.3 t) to a cis-lunar destination. SLS Block 2 CPL delivery mass to cis-lunar destinations should be in excess of 33,069 lb (15 t).
Note: All CPL delivery mass goals exclude associated PLA mass; therefore, this capability represents the mass of the actual payload delivered post separation from the EUS.
This section describes the SLS environments to which a spacecraft or payload will be exposed, both during ground processing and in flight.
Payload providers must perform an extensive, generalized series of system-level qualification and protoflight tests in order to achieve payload acceptance for spacecraft launched on the SLS. Such tests should be structured to encompass SLS block and payload configurations. Coordination with SPIE during test criteria development is encouraged to ensure process compatibility with SLSspecified operating regimes. To comply with NASA payload integration processes, payloads must demonstrate both design qualification and flight unit acceptance. Such criteria are assessed relative to payload structural integrity, along with operation failure modes with risk of impairing launch success when subjected to predicted SLS vibration, shock and acoustic loading environments. Qualification process reasoning and test data must be provided to NASA for secondary review to prevent mission goal incompatibilities and discrepancies between requirement definition and spacecraft verification. Furthermore, the intent is to ensure minimum standards regarding spacecraft build quality are satisfied. Protoflight procedures include testing of provided payloads at maximum tolerable SLS operating conditions to verify standards are met.