Artemis-I — formerly Orion / EM-1 (Exploration Mission-1)
In March 2018, NASA renamed the former Orion / EM-1 (Exploration Mission-1) to Artemis-I. Artemis-I will be the first integrated test of NASA’s deep space exploration systems: the Orion spacecraft, Space Launch System (SLS) rocket and the ground systems at Kennedy Space Center in Cape Canaveral, Florida. The first in a series of increasingly complex missions, Artemis-I will be an uncrewed flight test that will provide a foundation for human deep space exploration, and demonstrate our commitment and capability to extend human existence to the Moon and beyond. 1)
Orion EM-1, previously known as SLS-1 (Space Launch System-1) is NASA's first planned flight of the Space Launch System and the second uncrewed test flight of the Orion MPCV (Multi-Purpose Crew Vehicle). NASA, ESA, European and US Industry have teamed to develop the ORION spacecraft.
Under an agreement between NASA and ESA, ratified in December 2012, NASA’s new Orion vehicle for human space exploration missions includes the ESM (European Service Module), based upon the design and experience of ESA’s ATV (Automated Transfer Vehicle), the supply craft for the ISS (International Space Station).
ESA’s industrial prime contractor for ATV, Airbus Defence and Space of Bremen, Germany, is leading a European industrial consortium, developing this vehicle on behalf of ESA and working closely with NASA’s Orion US industrial prime contractor Lockheed Martin Space Systems.
Orion is the spacecraft that NASA intends to use to send humans and cargo into space beyond low earth orbit and to return them safely to Earth. It is being developed for crewed missions to cislunar space, asteroids, and then to Mars. The capsule is also planned as a backup vehicle for missions to the ISS (International Space Station). It will be launched by the NASA-developed SLS (Space Launch System) in 2019. 2) 3)
The first test flight of Orion was successfully completed in December 2014 with a Delta IV launch vehicle that launched the Orion Crew Module into a high elliptical orbit to demonstrate high-speed atmospheric Earth re-entry. This flight test did not include the ESM.
It is planned such that the Orion EM-1 flies a mission profile similar to what might be used in a future asteroid redirect mission. The un-crewed Orion travels to a lunar DRO (Distant Retrograde Orbit) mission (Figure 2), then returns on a trajectory calculated to achieve a high speed atmospheric entry on the order of ~11 km/s to demonstrate the performance and effectiveness of the Orion TPS (Thermal Protection System), as well as relevant environments prior to the first manned launch of the system. The total mission duration is 25 days (6 at destination). This will be the first flight of the ESM.
Orion EM-2 is planned as a Crewed High Lunar Orbit mission no later than the end of 2021. It is planned to spend several days in lunar orbit before performing a TEI (Trans-Earth Injection) burn to begin the return to Earth. The total mission duration is expected to be 14 days maximum (4 at destination).
Table 2: Overview of Orion ESM Design Reference Missions
Figure 1: Artist's rendition of the Orion EM-1 spacecraft in lunar orbit (image credit: NASA, ESA)
The Orion Vehicle
• CM (Crew Module)
• ESM (European Service Module)
• CMA (Crew Module Adapter)
• SA (Spacecraft Adapter)
• SAJ (Spacecraft Adapter Jettisoned Fairings)
• LAS (Launch Abort System)
The Service Module (SM) refers to the combined CMA + ESM + SA + SAJ.
The Crew Module and ESM (European Service Module), also referred to as SM, are physically interfacing via an interface ring called the CMA (Crew Module Adapter). The ESM is attached to the CMA for the duration of the mission. Just prior to the Earth’s orbit entry, the CMA separates from the CM for CMA/ESM disposal and the CM performs final reentry and landing operations.
NASA is responsible for development of the CM, CMA, SA, SAJ, and LAS elements of the Orion spacecraft. The CMA provides the structural, mechanical, electrical, and fluid interface between the CM and ESM. In addition, the CMA houses communication equipment, sublimators for thermal heat rejection, and power and data control/interface electronics. The SM is enclosed by three spacecraft adapter fairing panels (SAJ) which provide a partial load path from the CMA to SA but also protect the solar arrays, radiators, and thrusters from launch and ascent loads. The fairings will be jettisoned during the ascent phase or following main engine cut-off of the launch vehicle.
The SA (Spacecraft Adapter) provides the interface to the launch vehicle during launch. During launch and ascent, the ESM and SA will be enclosed by the SAJ. The SA attaches to the aft end of the ESM to the Launch Vehicle and includes the structural interface, separation mechanisms, and umbilical connectors for communication between the launch vehicle and the Orion Spacecraft. At launch vehicle burnout, the Orion Spacecraft separates from the SA at the ESM/SA separation plane.
Orion EM-1 CM (Crew Module)
The milestones for the Orion EM-1 mirror the path taken by the Orion EFT-1 (Exploration Flight Test-1) spacecraft. However, the Orion EM-1 will sport a number of improvements based on the experiences of the December 2014 test flight. NASA is beefing up the critical TPS (Thermal Protection System) that will protect astronauts from the searing heats experienced during reentry as the human rated vehicle plunges through the Earth’s atmosphere after returning from ambitious expeditions to the Moon and beyond.
Based in part on lessons learned from EFT-1, engineers are refining Orion’s heat shield to enhance the design, ease manufacturing procedures and significantly strengthen is heat resistant capabilities for the far more challenging space environments and missions that lie ahead later this decade and planned further out in the future as part of NASA’s agency-wide ‘Journey to Mars’ initiative to send humans to the Red Planet in the 2030s.
On all future flights starting with EM-1 (Exploration Mission-1), the Orion crew module must withstand the higher temperatures and speeds experienced during return from more distant destinations such as the Moon, near-Earth Asteroids and Mars. Orion's TPS is comprised of the 5 m diameter main heat shield covering the rounded base of the capsule as well as the grid of back shell tiles bonded around the vehicle from top to bottom. 7)
Starting with EM-1, engineers will bond Orion’s thermal protection system back shell tiles with a silver, metallic-based thermal control coating. The coating is designed to keep Orion’s back shell in a temperature range from approximately -65 to 290ºC prior to entry and afford protection against electrical surface charges in space and during reentry.
The pressure vessel is the primary structure of Orion’s crew module and is made of seven large aluminum pieces that must be welded together in detailed fashion. The first weld connected the tunnel to the forward bulkhead, which is at the top of the spacecraft and houses many of Orion’s critical systems, such as the parachutes that deploy during reentry. Orion’s tunnel, with a docking hatch, will allow crews to move between the crew module and other spacecraft. 8)
Orion’s pressure vessel is composed of seven large pieces of aluminum, three of which are the cone panels. The pressure vessel holds the atmosphere astronauts will breathe against the vacuum of deep space, forming the crew compartment. The three panels together form the angled mid-section around the crew module where the windows and hatch are located. While technicians have been joining other elements of the structure together since early September, the cone panels have presented a unique challenge for NASA and Lockheed Martin, the agency’s prime contractor for Orion. Engineers who have sought to reduce the crew module’s overall weight have encountered and overcome technical challenges. 9)
The lessons learned from the EFT-1 mission are being applied to the EM-1 (Exploration Mission-1) and EM-2 flight test configurations to optimize a system design that can smoothly transition into production.
Ten of the most important lessons from EFT-1 focused on the ability to evaluate the design with actual flight performance. This was critical for the remaining design process as NASA continues to develop the most advanced Human rated spacecraft ever built. The following are the top ten lessons learned from this test flight: 10)
1) Some deep space designs are classic for a reason: Every aspect of Orion’s design is driven by crew safety. Over 50 years of NASA’s investment learning the ins and outs of human spaceflight has provided direction for meeting deep space requirements. An example is LAS (Launch Abort System). About 6 minutes into launch, the LAS is jettisoned to save mass for the journey to deep space destinations. While other system configurations exist, design trade studies repeatedly highlighted the advantage of not carrying extra weight past the time it is needed. So because the Orion LAS supports missions to deep space, mass is king and any extra weight is an extreme hindrance to those missions. By shedding the Orion LAS early, this mass isn’t a burden to the vehicle for the entire mission, and other Orion systems are able to provide the abort function for the remainder of ascent to orbit.
2) Crew safety is built in, not bolted on: The EFT-1 tested and verified systems that are built “into the bones” of the spacecraft from the very beginning. This means not just the heat shield or the flight computers are designed for the rigors of deep space, but everything included in the fundamental systems and structure of the spacecraft are designed and built to specifications set forth by NASA. Every system design has a deep space requirement it is being built to right now: Orion’s seats are being designed to help prevent loss of consciousness as astronauts experience up to 5 G’s during high-velocity re-entries, the cooling system keeps the crew cabin about 25ºC despite its heat shield being heated to 2,200ºC during reentry, the built in stowage lockers are designed to double as a safe-haven during dangerous solar activity, the life support system allows for exercise since deep space missions require much longer stays in zero gravity, computers and avionics are designed to self-correct in case there is a failure and you’re months from home, and the crew module tiles are designed to protect from multiple micro-meteoroid strikes since the number of strikes will increase during missions that last months instead of days.
3) Reusability must be tied to reality: Part of flying for the first time in space is being able to make informed decisions about what we can realistically reuse following a deep space mission. After evaluating areas of water intrusion and corrosion, we’ve come to expect that many components in the crew module, especially inside the pressurized volume, or the hull where the crew sits, can be reused for later flights—components such as the computers, avionics and electrical distribution for example. The structure itself is more difficult to predict for reuse base on the unpredictable landing loads it might experience from a long journey from deep space with unpredictable landing sea conditions. The program is looking at options for reuse based on actual landing load data provided.
4) Designs mature as we Learn: Significant design changes are being made to optimize the EM vehicles. Thefuture design and analysis efforts will be simplified (Figure 5), overall mass has gone down, and recurring costs for production will be lowered. Establishing a common design philosophy has prevented a large number of engineering revisions and hardware changes. The Orion design is now incorporating lessons from EFT-1 and updated requirements for the crewed EM-2 mission.
5) Test like you fly to ensure success: After extensive testing on the ground in “Flight Like” scenarios EFT-1 successfully tested Orion in real flight environments, which could not be duplicated in ground simulations. The systems were subjected to the most critical crew safety requirement with the same rigors they will see when carrying humans. Systems verified included: thermal protection system, hardware separation events and the parachute system. This was a 100% mission success. 87 EFT-1 flight test objectives were identified in the early phases of the test development program. These objectives included verification of Orion’s subsystems ability to launch, control its trajectory with OFI (Operational Flight Instrumentation), complete all separation events, reenter the Earth’s atmosphere at 32, 200 km/hr and 2200ºC, land accurately in the Pacific Ocean and be recovered without damage. 81 of the 87 FTOs (Flight Test Objectives) were fully satisfied with six being partially met. Four of the seven FTO's were related to suspect DFI (Developmental Flight Instrumentation) performance. Two FTO’s related to the Crew Module Up-righting System were partially met, and one FTO related to structural measurements was partially met.
6) Organize for focus: Organizationally, Lockheed Martin and NASA agreed assigning a “mission director” to each test flight would allow program focus on the near term test milestones while maintaining a parallel program focus on the remaining DDT&E (Design, Development, Test, and Evaluation) efforts. EFT-1 was the first mission where this was implemented and proved to ensure the test was completed on time and within budget. This effort also saved over one year of development time for EM-1 by performing this in parallel versus chronologically.
7) Processes and reviews tailored: Since its inception the Orion program has spent significant time focusing on defining an efficient set of requirements to enable the design of Orion as well as optimize the concept of operations for a recurring program rhythm. In an effort to remain flexible the program continues to evolve the requirements based on lessons learned. On EFT-1 the program used a consolidated requirements document approach that was successful in limiting the number of requirements, individual requirements documents, and tiers of requirements documents. The team is using the success of this approach to streamline the EM requirements where possible. Requirements verifications were completed later than planned, partly because the verification work did not carry the same urgency as other launch campaign tasks. The team is looking at a more structured “waterfall” approach to prevent lower level verifications from conflicting with launch campaign activities.
8) Flexibility to accommodate change: Several program planning improvements were accomplished during the course of the program leading up to EFT-1. Improving the Orion IMS ( Integrated Master Schedule) by integrating the NASA “non-Prime” elements improved the continuity of the entire schedule past vehicle delivery. Through Monthly Orion Program Performance Review meetings, the assessments team continued to monitor the contractor performance. Data input for these reviews consisted of Integrated Master Schedule updates, CPRs (Cost Performance Reports), Financial Management Reports (533 inputs), contractor supplemental financial and schedule data, and through participation in the subsystem IPT (Integrated Product Team) meetings. The assessments team integrated these data each month to update the assessment and forecast for the EFT-1 launch date and financial position enabling program management to make rapid decisions.
9) Partnerships ensure Communication: After the President’s proposed cancellation of the Constellation program, NASA and Lockheed Martin recognized reductions needed to occur both on the government side and within Industry to keep the program intact. NASA identified all work within their scope as “Non-Prime” and all work within Industry as “Prime”. NASA and Industry were challenged with a $738.9 M reduction in funding in 2010 and needed to re-plan the entire program in anticipation of this reduction. The cost reduction initiatives they initiated have come to be known as “The streamlining of Orion” and have been used as examples of how NASA and Industry can work together to become more efficient and affordable. The continued emphasis on meeting program objectives within the government’s affordability range have resulted in a refined program plan that accomplishes all of its goals within the annual budget and without jeopardizing mission success. One example of this was the reduction of test flights from seven to four while maintaining the same requirements verifications.
10) Supply Chain must remain healthy: EFT-1 provided the Orion team the ability to exercise 80% of the supply chain that will be utilized for future production vehicles. This includes subcontractors currently spread across 42 states with 60,000 parts being received by NASA or its prime subcontractor in 3 major integration facilities. Throughout this endeavor the processes were validated, enhancements were made and opportunities for improvement identified. EFT-1 helped establish 6 key principles needed to ensure effective supply chain management:
• Centralized supply chain ownership/management offers advantages over IPT ownership
• Need for timely actionable information, including stable engineering design documentation, is mandatory
• Consistent, effective and constant communication is mandatory
• Management of the supplier certification and work load
• Manage with the tools, don’t expect the tools to manage
• Work within the system.
In summary, NASA and Lockheed Martin are taking what we have learned from EFT-1 and years of government investments to make improvements, fly again, making improvements again, and developing a spacecraft that we can proudly stand behind and say we are confident in its abilities to take humans into deep space and bring them safely home.
Some examples of Orion design changes as a result of lessons learned from EFT-1 deals with the structure of the EM.
The Orion design for future Exploration Missions experienced many design optimizations as a result of the EFT-1 flight results. The following section highlights several changes to exemplify optimizations that was realized. Note: (For more information, the reader is referred to Ref. 10).
Structures: Reduced parts and weld assemblies required on the pressure vessel include:
- Reduction from 6 to 3 cone panels, and reduction of cone section welds from 12 to 3.
- Reduction from a 3 piece welded aft bulkhead design to a single spun-formed design.
- Change to an “Apollo-gusset” design in the aft-bay to help reduce the number of separable-parts, and combine structural elements (reduced thruster pod support structures and harness support structures, etc.).
Figure 6: A simplification and reduction of structural parts has been realized by the lessons from EFT-1 (image credit: NASA, Lockheed Martin)
Propulsion: Tubing subassemblies on EFT-1 were custom fit for propulsion system integration (e.g. fit-up tube assembly, cut, face, install heaters / temp sensors, reinstall, repeat as required). On EM-1 no trim to fit will be required on the vehicle. Tubing subassemblies will be delivered in their net shape and subassembly welds will be proof & leak tested prior to delivery.
PCAs (Pressure Control Assemblies) on EFT-1 required a custom fit for propulsion system integration (similar to Tubing subassemblies). On EM-1 there will be improved packaging and attachment methods including PCA welds proof & leak checked prior to delivery.
Propellant and Pressurant Tanks on EFT-1 were difficult to integrate and perform welding operations. The EM-1 design created a new inlet / outlet orientation to ease integration and welding operations.
RCS (Reaction Control System) thrusters on EFT-1 had support struts that have been simplified to a Pod design and attachment method eliminating all struts.
Figure 7: Left: EFT-1 Roll Left Thruster, Right: EM-1 Roll Left Thruster Pods (image credit: NASA, Lockheed Martin)
TPS (Thermal Protection Systems): The Heatshield Avcoat design improvement included changing from a monolithic, individual cell injection process to a “Block Avcoat” design where the heat shield blocks could be automated in production while increasing material properties performance. The use of thermal tape that was used on the EFT-1 heatshield (and similar to Apollo) was changed to use on all backshell and heatshield surfaces for all future Exploration Missions.
Micrometeoroid and Orbital Debris Performance: There are a number of lessons learned in the realm of MMOD (Micrometeoroid and Orbital Debris) environments and analyses, which are being examined and used to influence change for the EM and beyond missions. Chief among these lessons learned is that time spent in the high orbital debris flux altitude band, approximately 650 km to 1200 km, should be minimized. The MMOD analysis performed on EFT-1 was the first such analysis of a human-rated spacecraft above ISS ( International Space Station) altitudes since man-made orbital debris has become an issue, and found the risk at these altitudes (600 km – 1600 km) to be much higher than anticipated; Apollo was too early in spaceflight history for much OD (Orbital Debris) to have accumulated yet. Post-flight inspection of the EFT-1 capsule indicated that actual MMOD exposure may have been even higher than the analyses had assessed. Based on this lesson learned, the parking orbit of EM-1 has been reduced from 3.5 hours in a high-MMOD-risk orbit (2 orbits), to half that time (limited to 1 orbit) to reduce MMOD risk. And as EM-2 trajectory trades and analysis is ongoing, avoiding these altitudes has been accepted as a prerequisite in the EM-2 trajectories. Largely anchored on this finding, other lesson learned recommendations include: increasing MMOD protection in upper-stage and other critical hardware, re-assessing window damage cause and remediation, and continued post-flight inspections and MMOD environment recommended updates based on EM flights.
Radiation Summary: Orion is the first spacecraft that addresses crew radiation protection as an integral part of the vehicle design. We are using a state-of-the-art radiation analysis process to analyze the shielding provided by the vehicle and quantify the internal radiation environment. This analysis is based on the full fidelity Orion CAD model and space radiation environment models. Due to the EFT-1 trajectory passing through the very core of the van Allen proton belts, intravehicular radiation environment was equivalent to 4-6 weeks inside ISS (in terms of cumulative exposure). Thus EFT-1 presented a valuable opportunity to validate our analysis procedure and ultimately improve crew radiation protection for future manned missions. This opportunity was materialized by flying six RAMs (Radiation Area Monitors) on EFT-1. Radiation measurements were in very good agreement with pre-flight predictions, confirming the validity of the radiation analysis approach and providing confidence in our efforts to maintain crew radiation exposure ALARA (As Low as Reasonably Achievable) consistent with NASA requirements.
An important lesson learned for electronic components radiation hardening refers to the importance of considering time variability of the environments in the Single Event Effect rate calculations. The EFT-1 proton environment varied dramatically throughout the mission. Accounting for this time variation was an essential component in selecting the appropriate level of redundancy in critical systems such as the VMC/FCMs (Vehicle Management Computer / Flight Control Modules). This lesson continues to apply to Exploration Missions. Stochastically occurringSPEs ( Solar Particle Events) may cause significant temporary increase in the radiation environment, and critical systems need to design appropriately. A related lesson learned reflects the importance of assessing timing of critical mission events with respect to the radiation environment. Environment assessments performed for EFT-1 drove modification of the mission timeline such that Avionics intensive mission events be executed outside of the core of the van Allen belts. This too is a lesson learned relevant for future missions especially in off-nominal conditions that may expose the spacecraft to high radiation environments.
EM-2 (Exploration Mission-2): EM-2 will be a crewed test flight that will be launched on an SLS and enter lunar orbit in 2022 to verify the capability of Orion to successfully launch, perform a crewed mission, and return them safely to Earth. Figure 2 depicts one option for a lunar mission. The EM-2 configuration will consist of:
• A fully functional launch abort system
• A fully functional Crew Module that includes a crew of up to four and all Cis-Lunar life support systems
• A fully functional Service Module.
Figure 8: EM-2 Design Reference Mission (image credit: NASA)
Figure 9: Technicians with Lockheed Martin, NASA’s prime contractor for Orion, are welding together the pieces of the spacecraft's pressure vessel at Michoud Assembly Center in New Orleans, LA (image credit: NASA)
ESM (European Service Module)
The ESM implements four major system functions to Orion (Ref. 2):
• provides thrust for orbital maneuvers and attitude control after upper stage/launch vehicle separation
• generates electrical power and distributes it to the ESM users and to the CM/CMA
• regulates heat for the life support and avionics equipment during the orbital phases of the mission
• stores and provides to the CMA/CM potable water, oxygen, and nitrogen.
In addition, it ensures structural spacecraft integrity during launch and in-orbit maneuvers. The ESM can also provide additional volume and other resources on select missions for accommodating science, engineering demonstrations, development test objectives, and deployment of lunar infrastructure equipment during the cruise and lunar orbit phases of lunar missions. This volume provides electrical power distribution, network access for command and control interfaces, and structures and mechanisms.
The architecture of the module has been developed based on the ATV spacecraft concept (five successful missions to the ISS),modified to cope with the different mission requirements and the man-rating approach for beyond LEO missions. 11)
Figure 10: Illustration of an ATV cargo freighter in flight with deployed solar arrays (image credit: ESA)
The resulting ESM architecture is depicted in Figure 11, for comparison. The architecture of the system and subsystems hereafter proposed is the reference design for the Lunar Sortie Mission. Changes in the architecture are expected for the other missions to tailor the configuration to mission needs, to remove unnecessary HW and to optimize the launch mass.
European Industrial Consortium: ESA entrusted the development of the ESM (European Service Module) to a consortium of European industries led by Airbus Defence and Space. The consortium of companies was selected to reuse the experience and industrial heritage of the very successful ATV (Automated Transfer Vehicle). Airbus Defence DS Germany as the Prime Contractor is responsible for all system-related work. This includes: 12)
- The overall Management of the contract
- The overall System Engineering activities
- The Management of the procurement activities
- The System Product Assurance and Safety activities
- The liaison with Lockheed-Martin (NASA prime contractor of the Orion vehicle).
Airbus DS, France is responsible for part of the system engineering, ground software, Helium pressurant tanks and simulation facility.
On Subsystem level, Airbus DS, Germany is responsible for the System Engineering, the Propulsion, Power and Avionics Subsystem Engineering as well as for GSE and AIT activities. The responsibility for the development of the other sub-systems or equipment has been distributed among European companies as follows:
TAS (Thales Alenia Space) Italy: the Structure, Thermal and Consumable Storage Subsystems, supported by the following level-2 subcontractors:
- RUAG, Switzerland: secondary structures
- SONACA, Belgium: tank bulkhead
- APCO, Switzerland: MDPS (Meteoroid and Debris Protection System)
- CRISA, Spain, TCU (Thermal Control Unit)
- Prototech, Norway: nitrogen filters
- MEWASA, Switzerland, Water tank bellows.
Dutch Space, The Netherlands: the SAW (Solar Array Wings), supported by the following level-2 subcontractors:
- SELEX SE, Italy for the Photo Voltaic Assemblies
- RUAG, Switzerland: deployment dampers.
Airbus DS, Germany: the propellant tanks, Propulsion Drive Electronic and reaction control thrusters.
RUAG, Switzerland: the SADA (Solar Array Drive Assembly) composed of the mechanism and electronic unit.
SELEX SE, Italy: the PCDU (Power Conditioning and Distribution Unit)
Thales Alenia Space, Belgium: PRU (Pressure Regulation Unit)
Antwerp Space, Belgium: the Electrical Ground Support Equipment Front Ends, supported by the following level-2 subcontractors :
- Clemessy, France
- Rovsing, Denmark.
APCO, Switzerland: Mechanical Ground Support Equipment.
Latelec, France: avionics and power harness
TESAT Germany and Alter, Spain: CPP (Centralized Parts Procurement) scheme for EEE (Electrical, Electronic and Electromechanical)-parts.
Equipment suppliers encompasses Vacco (US), Moog (US), Sofrance (F), Cobham (US), Vivace (US).
Physical architecture: The ESM is a cylindrical unpressurized module which interfaces at its bottom to the SA and at its top to the CMA. The OMS-E (Orbital Maneuvering System -Engine), i.e. the main engine, protrudes into the SA. Equipment on the tank platform is also allowed to protrude into the CMA. The total height of the ESM is 4.0 m.
The cylindrical shape is defined by the external radiators which enclose the body of the ESM. The function of the radiators is twofold, i.e. to radiate heat, and to serve as the first barrier of the MDPS (Micrometeoroid and Debris Protection Subsystem ). Mounted to the back side of the radiators are Nextel and Kevlar blankets which serve as the second barrier. The externally mounted RCS (Reaction Control System) pods and SAWs (Solar Array Wings) must be designed such that these respect the Orion SAJ allowable envelope.
The internal accommodation of the ESM subsystem equipment is highly dependent to the primary structure architecture, based on seven separate bays for accommodation, six located circumferentially about the central bay. The pressurant tanks are accommodated in the center bay. The four largest bays are used for the accommodation of the propellant tanks which dominate the available volume. The remaining volume provides accommodation to the avionics equipment, CSS (Consumable Storage Subsystem) water tanks, harnessing and tubing.
The top surface of the ESM is dominated by the protruding tank domes for the propellant subsystems, one of the helium pressure tanks, the four tanks for Oxygen and Nitrogen, and the Flow Control Assembly for the active cooling subsystem. These elements protrude into the CMA, while maintaining a specific minimum distance to the CM Heat Shield. For these elements above the tank platform, the CMA provides the MDPS protection.
The lower surface is dominated by the OMS-E, which serves as the main engine of the ESM, and the eight auxiliary thrusters. Within the lower platform, a panel is incorporated through which the UPC (Unpressurized Cargo) is installed, and optionally ejected.
Table 3: ESM characteristics for Lunar Missions
Functional architecture: Orion functionalities are generally shared between ESM and CMA/CM. Functional Chains gather equipment of the ESM which participate to a same service provided by Orion for the accomplishment of its mission.
Figure 12 depicts the overall avionics subsystem architecture and interfaces. The interface to the SLS launcher is depicted on the left side of the drawing, and the CM and CMA (labeled as SM-CM I/F Adapter) are on the right side. The main link with Orion CM on-board computers is the ODN (On-board Data Network), based on a time-triggered Ethernet solution. This interface ensures the connection of all ESM avionics to the on-board computers and allows the ESM to receive commands and to deliver monitors.
Few discrete lines ensure the independent command of ESM power subsystem electronics, allow to acquire signals from 8 sun sensors used as back-up AOCS sensors and allow the transfer of data acquired from the development flight instrumentation.
The four SAWs (Solar Array Wings) ensure a maximum electrical power production of 11.2 kW, to cope with a CMA/CM power demand of maximum 7.3 kW to be respected with one wing failed.
Four power buses allow powering to / from the CMA/CM depending on mission phases. Additional and independent power lines are provided for the wireless cameras installed on top of SAWs.
Finally, a dedicated pass-through harness connects the CMA to the SA/SAJ and launcher to allow transfer of monitoring data and separation commands.
The ESM provides translational and 3-axis attitude control for the spacecraft, stores consumables for the crew module and provides power via the solar arrays. The ESM is also being designed to carry cargo. The main subsystems of the ESM are: 13)
• Structure Subsystem
• TCS (Thermal Control Subsystem)
• CSS (Consumable Storage Subsystem)
• PSS (Propulsion Subsystem)
• EPS (Electrical Power Subsystem)
• Avionics Subsystem
Figure 13: Top view of the ESM and its elements (image credit: ESA, Airbus DS)
Figure 14: Illustration of the ESM architecture and principal layout of main elements (image credit: ESA, Airbus DS)
Structure subsystem: The ESM astructure consists of the following major elements:
• Primary structure
• Secondary structure
• MDPS (Meteoroid and Debris Protection).
The primary structure transmits the SLS launch loads to the upper composite composed of the CMA, the CM, and the LAS, then supports the upper composite aerodynamic and inertial loads as well as the ESM inertial loads generated by the ESM internal equipment. The design is based on a shared load path between the ESM and the external SAJ fairings. The objective is to minimize the loads applied on the ESM mechanical structure in order to optimize its mass which is propelled all along the mission, whilst the SAJ capability is sized to the extent possible by the main launch phase, since the SAJ mass is jettisoned early during the launch. The primary structure is composed of:
• 6 longerons (machined aluminum) linked to the CMA frame and six pyronuts that separate the ESM from the Spacecraft Adapter
• Tank bulkhead (machined aluminum) supporting the four propellant tanks, as well as the gas delivery CSS tanks, ensuring the main link between the CMA lower interface ring and the rest of ESM structure.
• "Radial" shear webs and internal "square" webs assembly (composite sandwich panels), housing most of the ESM equipment, including water tanks and SAW support frames and propulsion high pressure system. It forms the ESM core attached to the tank bulkhead, longerons and lower closeout panels. The main engine is attached to the central square tube panels via struts transmitting the thrust.
• Lower Platform (machined aluminum) on which the equipment is connected: OMS-E main engine, auxiliary thrusters supports, RCS pods supports, SAD (Solar Array Drive) mechanism, and PIE (Propulsion Isolation Equipment).
• MDPS covers and aft closure panels to protect the ESM from the MMOD (Micrometeoroid and Orbital Debris) environment.
The secondary structures support the ESM equipment carrying the inertial and dynamic loads of the equipment during launch and transmitting in-orbit thrust and inertial loads.
The MDPS is partly metallic and partly composed of Nextel reinforced MLI (Multi-Layered Insulation). Debris and meteoroid particles with velocities up to 24 km/s hit the outer wall, forming a cloud of lower energy particles which are then contained by the inner wall, preventing any penetration of the ESM.
Figure 15: Primary and secondary structures (image credit: ESA, Airbus DS)
TCS (Thermal Control Subsystem): The TCS includes an ATCS (Active Thermal Control Subystem), a PTCS (Passive thermal control Subsystem) and the TCU (Thermal Control Unit).
11) ATCS (Active Thermal Control Subystem): The ATCS is designed to collect the thermal loads from the CM and from the ESM powered equipment and to reject them toward the space radiative sink. The ATCS architecture is based on a Single-phase Fluid Loop Architecture using the HFE-7200 coolant to collect and to transfer the heat loads from both ESM avionics (via cold plates) and CM (via Inter-Loop Heat Exchanger) and to reject them through specific body mounted radiators. The ATCS is composed of two fully independent loops working simultaneously (hot redundancy approach). The schematic of the ATCS architecture is presented in Figure 16.
Each ATCS loop is composed by the following main components:
• 1 Radiator Assembly commonly used for both loops and composed of 4 full height radiators and 2 split radiators (with lower part not removable) mounted in serial configuration
• 1 CP (Cold Plate)
• 2 online APSBs (Absolute Pressure Sensor Blocks)
• 1 wet on-line TSB (Temperature Sensor Block)
• 1 FCA (Fluid Control Assembly) including redundant pumps with relevant passive accumulator and a TWMV (Three-Ways Modulating Valve)
• Hard/flex hoses, tees, couplings and restrictors.
The FCA is controlled by the TCU to provide constant mass flow rate inside the loop. In particular two flow rate set points are foreseen, one for nominal operation and one for contingency (with only one loop operative). The FCA also includes the TWMV, controlling the temperature at the ESM-CMA interface.
The external surface of the radiator panels is coated with a specific paint characterized by the following thermal-optical characteristics:
• Alpha (α) = 0.2
• Emissivity (ε) = 0.8.
The Cold Plates are devoted to collect the entire thermal load from the ESM avionics boxes. Their design is constituted by a stainless steel channelling enclosed in an aluminum casting acting as a plate in direct contact with the thermally active unit. Different Cold plate sizes, to satisfy the different configuration needs, are foreseen.
Figure 17: Photo of a sample radiator panel (image credit: ESA, TAS)
12) PTCS (Passive Thermal Control Subsystem): The PTCS provides the thermal control of ESM hardware (propulsion, CSS, power and avionic items) and reduces the temperature gradients and minimize heat flows through the internal elements. The PTCS has two main components, heaters (with thermistors and wire heater) and insulation (MLI thermal blankets, including some specific high temperature MLI blankets, for thrusters and engine nozzle thermal impingement protection).
The MLI protects the internal parts against the external environment and heater lines compensate the heat leaks toward space. Different MLI typologies, in terms of composition, have been identified for the different applications on the ESM. The MLI composition mainly depends on exposure or not to the space, geometry & fixation interfaces and exposure to thrusters plume flux and nozzle radiation.
Two cold redundant heater lines managed by the TCU assure thermal control and temperature uniformity in the Orion internals and for the local thermal control of specific items. In addition, heater lines managed by thermostats and powered by the PCDU (Power Control and Distribution Unit) provide a further redundancy to the environmental control function.
13) TCU (Thermal Control Unit): The TCU is designed to ensure the management of the TCS and of the CSS.
For the ATCS, the TCU acquires the various parameters from the ATCS sensors, monitors and commands the ATCS valves, and issues commands to the FCA (Flow Control Assembly).
For the PTCS, the TCU commands the heater chains and monitors the thermistors. It operates according to activation and de-activation thresholds defined and modified by the CM on-board computers, while ensuring bus interface and power supply logic.
For the CSS (Consumables Storage Subsystem), the TCU commands the valves and monitors their position. It also monitors the pressure and temperature parameters in the subsystem to support the system FDIR (Failure Detection, Isolation and Recovery).
Figure 18: Layout of the TCS (Thermal Control Subsystem), image credit: ESA, Airbus DS
CSS (Consumable Storage Subsystem): The CSS provides potable water, oxygen and nitrogen to the CM. It can also provide water for the sublimator located in the CMA via a dedicated kit composed by its own tank assy and distribution system, to be installed optionally for specific missions. It consists of the following major elements:
• WDS (Water Delivery Subystem)
• GDS (Gas Delivery Subsystem)
The architecture of the WDS is composed of the following components:
• Water tanks (metal bellow technology)
• Water on/off valves, (isolation valves)
• Temperature and quantity sensors
• 2 distribution lines towards the CMA I/F to supply water for the different CM/CMA needs.
The architecture of the GDS is composed of the following components:
• Gas tanks
• Gas pressure regulators
• Gas on/off valves, (isolation valves)
• Gas relief valves
• Temperature and pressure sensors
• Hydrophobic filters upstream the water tanks
• Distribution lines towards the CMA I/F to supply oxygen and nitrogen
• Distribution lines towards CMA I/F for GDS filling on ground from the GSE.
In the case of a configuration with WSK (Water Sublimator Kit) a set of 3 additional water tanks is introduced and interconnected to the others tanks and lines of CSS.
Figure 19: Layout of the CSS (image credit: ESA, Airbus DS)
PSS (Propulsion Subsystem): The propulsion subsystem design has to cope with a complex mission scenario. It includes three different types of engines / thrusters. Each of these thrusters utilizes the same propellants: MON-3 (Mixed Oxides of Nitrogen) as oxidizer and MMH (Monomethylhydrazine) as fuel.
The single Main Engine, which is one of the Shuttle OMS-E engines (27.7 kN), is delivered by NASA as GFE (Government Furnished Equipment) and will be used during ascent abort and trans-earth injection maneuvers as well as orbit change maneuvers. This engine is gimballed with an amplitude of ± 7° around both axes (pitch and yaw). The actuation is performed by the TVC (Thrust Vector Control).
The 8 auxiliary thrusters which are similar to those used on ATV (490 N) are used during ascent abort and launcher separation together with the main engine. They are also used as backup to the main engine for trans-earth injection and for trajectory correction maneuvers.
The 24 RCS thrusters which are the same as used on ATV (220 N) provide the impulse necessary for attitude control, small maneuvers and forced translation during Rendezvous and Proximity operation. These thrusters are accommodated into six pods. Two pods are composed of four roll thrusters, placed along the ESM lateral surface under the CMA, while four other pods are composed of four thrusters, placed along the ESM lateral surface under the level of the first four pods. Each propellant type is stored in two propellant tanks, both sets of propellant tanks being arranged in a series configuration.
The propellant distribution comprises several electromechanical valves for propellant isolation and pressure transducers interconnected via pipes. The assembly ensures isolation of the engine / thrusters from the propellant tanks during launch, docked phase to the ISS and other mission events where it becomes necessary to isolate the engine / thrusters from the propellant storage. The propellant distribution network provides the connections between the propellant tanks and the isolation valves and then onwards to the thruster assemblies. Fill and Drain Valves and the related access lines are used during ground testing and for loading and unloading of the propellant and pressurized tanks. They are located in the CMA to enable access even after the SAJ around the service module has been installed.
The propulsion subsystem is controlled by the PDE (Propulsion Drive Electronics) which handles all nominal propulsion related commands issued by the CM and provides feedback via the CMA PDUs (Power Distribution Units) and the PDUs to the vehicle management computer interface and the ESM network. The PDE controls all thrusters, the TVC and the associated valves, instrumentation and leakage detection.
The PRU (Pressure Regulation Unit) controls and monitors the operation of the pressure regulation valves, based on set points defined and modified by the on-board computers. The EPR (Electronic Pressure Regulation) selected for ESM has the advantage to optimize the pressure of propellant and oxidizer according to the engine in operation, at the price of an increased development risk for this function.
The EPR of the propulsion system propellant is a new technology for the European industry. The ATV design featured a mechanical pressure regulation like the propulsion system of the Shuttle OMS-E engine. The electronic pressure regulation allows a better control of the pressure, hence the fuel and oxidizer consumption is enabling to reduce propellant margins and the overall mass at launch. Airbus DS decided to venture in that new technology as a result of a trade-off taking into account the positive results obtained by Lockheed-Martin in breadboard testing. However, in the development of a new technology there is always the risk of not being successful especially if the development time is constrained.
To mitigate the impact to the project in case of unsuccessful development, it was decided to ensure that reverting to mechanical pressure regulation would be possible with minimum schedule impact. NASA provided Shuttle mechanical pressure regulators and valves to VACCO, the original manufacturer and Airbus DS ordered tests to ensure the hardware would satisfy the performance required for the ESM. Breadboard testing of the electronic regulation performed in Europe up to now delivered promising results and the probability of having to revert to mechanical pressure regulation is becoming more remote.
Figure 20: Layout of the propulsion subsystem (image credit: ESA, Airbus DS)
EPS (Electrical Propulsion Subsystem): The EPS has the function to generate the power for all modules of Orion. It manages the power provided by its 4 SAW (Solar Array Wings). The PCDU (Power Control and Distribution Unit) provides the power I/F to the SAW and CMA, distributes electrical power to ESM users and protects the power lines.
The power generation part of the ESM EPS consists of four SAW units. Each wing is composed of three deployable CFRP (Carbon Fiber Reinforced Polymer) rigid panels covered with triple junction GaAs solar cells forming nine (9x) sections of solar cell strings. In nominal condition the SAW can supply a total 11.2 kW. Each SAW is linked to the ESM structure by a two degree of freedom SADA (Solar Array Drive Assembly). The SADA ensures the power and signal transfer from the SAW to the PCDU. It is composed of SADM (Solar Array Drive Mechanism) and a (SADE (Solar Array Drive). The SADM allows orienting the SAW in two independent axes. In the Sun-tracking mode the inner axis can swivel between -35º and +25º, while the outer axis has a continuous rotation capability (0º; +360º).
SADM of the ESM has a unique design incorporating a two axis gimbal. An inner axis provides rotation of the SAW about an axis perpendicular to the ESM longitudinal axis, and an outer axis which rotates the SAW about its own longitudinal axis. The two-axis capability is necessary for two reasons:
1) allow maximum sun tracking to meet the power requirements for particular vehicle attitudes of certain mission phases. Insufficient power is generated by the SAW with a single (roll) axis SADM not providing the avoidance capability of the shadowing effect of both the Orion vehicle on the SAW and the SAWs on each other
2) insure the structural integrity of the SAWs under injection maneuvers. For the trans-lunar injection performed with the upper stage of the Space Launch System rocket, actually the iCPS (interim Cryogenic Propulsion Stage), the arrays are canted backwards to sustain in deployed configuration the 1 g acceleration load. For the trans-earth injection performed with the ESM main engine, the acceleration is less severe and the arrays have to be canted forwards to prevent damages from the OMS-E engine plume while minimizing the load on the SAWs.
The SADM also allows repositioning (canting) of the SAW to reduce the loads on the SAW and SADM during the different Orion orbital maneuvers. At TLI (Trans-Lunar Injection), a 0.5 g acceleration is generated by the iCPS engine and at LOI (Lunar Orbit Insertion) and TEI (Trans-Earth Injection) a 0.3 g deceleration is generated by the OMS-E engine.
As shown in Figure 21, in order to reduce these loads to an acceptable level, the SAWs need to be repositioned by the SADM. In TLI, the SAWs are thus canted during the maneuver to inner axis = - 60º / outer axis = 0 º. In LOI/TEI, the SAWs are canted to inner axis = +55º / outer axis = 0º.
The EPS distributes the generated power to the CMA/CM through four independent 120 V busses. The power interfaces of the PCDU protect each unregulated bus from overload or short circuit failures on feeder busses outside the PCDU. They are implemented with power devices connected in parallel and independently operated for power flow from PCDU to CMA PDU, plus a number of power diodes for power flow from CMA PDU to PCDU.
Avionics Subsystem: The high-level management of the vehicle and its functions is performed at the level of the Orion on-board computers. The ESM is controlled and monitored by specific electronic controller units. All units are connected via the ODN (Onboard Data Network) to the CMA/CM. All vehicle control and management software resides in the CM computers. The ODN is a three plane TTGbE (Time Triggered Gigabit Ethernet),1000Base-CX, according to standard SAE AS6802.
The Time-Triggered Gigabit Ethernet ODN provides a set of time-triggered services implemented on top of standard IEEE802.3 Ethernet. It provides the capability for deterministic, synchronous, and congestion-free communication, unaffected by any asynchronous Ethernet traffic load. By implementing this standard in network devices (network switches and network interface cards), Ethernet becomes a deterministic network. This means that distributed applications with mixed time-criticality requirements (e.g., real-time command and control, audio, video, voice, data) can be integrated and coexist on one Ethernet network.
Each of the ESM controller units is connected via a SNIC (Standard integrity Network-Interface Card), according to Honeywell ICD8273722, to the LAN switch located in the CMA. The SNICs have three ports for two failure-tolerant communication on three network redundancy planes.
Development: The challenging schedule of the program, with the ESM delivery planned for early 2017 and the EM-1 launch planned for 2018, is influencing the way the ESM is verified. Qualification has often to run in parallel to EM-1 flight model manufacturing and integration, bringing risk to the program. This risk is carefully monitored and risk reduction strategies are in place to ensure the program can proceed smoothly. The ESM verification is mostly accomplished at module level. A limited number of verifications is planned on the EM-1 flight model.
- The system PDR (Preliminary Design Review) was successfully concluded in the summer 2014.
- The system CDR (Critical Design Review) is scheduled for the end of2015 and will conclude with the board in February 2016.
Table 4: Summary of key aspects for an international partnership in the Orion program