As the on-going political refinements to the FY2011 budget proposal raise the hope of utilizing “Shuttle’s Legacy” in a Heavy Lift Launch Vehicle (HLV), the third article – based on the findings of the expansive SD (Shuttle Derived) HLV assessment presentation – outlines both the use of the HLV in returning to the Moon, and the joint role of working with commercial and international vehicles in a Beyond Earth Orbit (BEO) architecture.
As covered in the two previous articles, a final assessment of the SD HLV options (726 page presentation, available on L2) – namely side mount and inline vehicles – provided expansive information on the development, costs, timeline and ability of the vehicle that would be a natural successor to the Space Transportation System (STS – Shuttle).
Immediate mission requirements, such as logistical resupply of the International Space Station (ISS), is catered for in the FY2011 budget proposal via yet-to-launch commercial vehicles, in tandem with overseas assets such as ATV, HTV and Progress, ahead of manned requirements to take over from purchasing Soyuz seats for US crewmembers during the upcoming gap.
However, the future outlined in the budget proposal only makes a passing reference to an eventual goal of a manned mission to Mars, sometime in the 2030s. It also cancels the Constellation approach to returning to the Moon.
While the SD HLV assessment is not a policy document – instead finalizing all the evaluations into the SD HLV, potentially for the benefit of documenting the work for a commercial entity picking up on the design – the findings do provide interesting architecture approaches for both Lunar and BEO missions.
SD HLV Lunar:
Utilizing the roadmap of the Program Of Record (POR) vehicles – such as the Orion manned vehicle, and Altair Lunar Lander – the assessment morphs the baselined HLV into providing the launch vehicle role, as opposed to the POR’s Ares I and Ares V’s 1.5 architecture, whilst trading the Earth Orbit Rendezvous (EOR) and Lunar Orbit Rendezvous (LOR) approaches.
“The human lunar mission in particular was a focus of the study team in matching HLV capabilities to lunar mission requirements. The NASA Orion CEV and Altair Lunar Lander designs were utilized as element requirements for such missions,” opened the Lunar section of the presentation.
“The study also examined approaches that retained the same mission requirements, but opted to conduct such missions in alternate scenarios more suited to HLV than the Ares I + Ares V ‘1.5 Launch Architecture’. Two operational approaches include Earth Orbit Rendezvous (EOR) and Lunar Orbit Rendezvous (LOR).
“Figure 5-9 (click image for larger screenshot) shows the performance trades for the two operational scenarios for the element (CEV and lander) mass requirements and the HLV Trans-Lunar Injection (TLI) capabilities for upper stage designs that use the rocket engines noted. The operational scenarios are described in the next sections.
“In all cases the study concluded that two Block II HLV launches are sufficient to perform a lunar mission that not only exceeds Apollo capabilities by a wide margin, but also captures the main requirements of the existing NASA lunar mission. Appendix L provides the details of the feasibility assessment for the lunar design reference mission.”
The large Appendix L section of the presentation shows two SD HLVs, one unmanned with Altair, one manned with Orion, both with Earth Departure Stages (EDS), providing pros and cons of EOR and LOR mission profiles – with the latter given preference. Both profiles are summarized in the Lunar section of the presentation.
“The first Block II HLV launch places an EDS stage into orbit along with the lunar lander. With suborbital staging the mass in LEO is maximized, with the EDS retains part of its propellant load for the TLI burn. The second launch occurs within days of the first and places the CEV and an EDS stage into a rendezvous orbit that allows for subsequent docking of the elements in a series configuration.
“The first TLI burn exhausts the propellants in the first EDS stage that is then jettisoned. The second EDS stage performs the remainder of the TLI delta-V burn requirements. The lunar lander performs the Lunar Orbit Insertion (LOI) burn to place the combined CEV and lander into low lunar orbit.”
The staging effect provides for increased TLI performance in the absence of a larger, single EDS stage comparable in size to one launched by Ares V.
The same profile outline was given for the LOR Lunar Sortie mission, again utilizing two launches of the HLV.
“Figure 5-12 shows how two HLV launches place the lunar lander and CEV into orbit separately and how the attached EDS stages then are used to send the individual elements into lunar space. In the LOR scenario shown, the lunar lander and CEV each perform their own LOI burns to place themselves into a low lunar orbit.
“The two elements then perform a rendezvous and docking. Figure 5-9 shows that the Block II HLV launches have sufficient performance to conduct this LOR mission with the CEV and lunar lander elements redesigned for LOR.”
Based on ten key factors of the mission profile, LOR was deemed as preferable on eight elements, compared to EOR’s two – namely Post Trans Lunar Injection (TLI) lifeboat and Lunar Orbit Insertion (LOI) burn. The twin HLV architecture also provides benefits over the POR’s 1.5 Ares I and Ares V approach.
“From this the HLV team concluded that LOR was the preferred operational scenario for a human lunar mission,” noted the presentation. “An alternative LOR scenario would have HLVs deliver high performance EDS stages that are sized to also perform the Lunar Orbit Insertion (LOI) burns to deliver the Orion class CEV or a lunar lander to lunar orbit. Once in lunar orbit, the CEV and lander would perform a lunar orbit rendezvous as before.
“This twin HLV launch LOR architecture offers advantages over the EOR architecture employed by the Ares I/Ares V systems by 1) making the lunar lander smaller by removing the LOI burn requirement from the descent stage, 2) minimizing the propellant boil off in the EDS while parked in Earth orbit, and 3) opening up a 2-month launch window for the CEV once the lander is safely in lunar orbit.
“Several propellants were evaluated for the lunar lander. LOX/LCH4 is presented based on its Mars forward, space storability, high density, and performance characteristics. The use of LOX/Methane on the lander allows the airlock to be package down low for ease of crew access to the lunar surface.
“This also allows the lander to be packaged inside a smaller HLV shroud of less than 7.5 m, which reduces shroud mass and aerodynamic drag, thereby improving launch mass.”
SD HLV BEO – Propellant Depots:
Commercial vehicle additions to the HLV approach – by way of the Upper Stage – are focused on in the BEO mission options, noted as higher energy missions. In providing benefits to Lunar and Deep Space missions, the assessment notes the addition of propellant deports would significantly improve the overall exploration architecture.
“HLV capabilities for higher energy missions beyond LEO were examined. This required the use of upper stages. Existing stages from Delta IV and Atlas V and new stage designs using RL-10B2 (existing), RL-60 and J-2X LOX/LH2 rocket engines were used estimate payloads to Geosynchronous Transfer Orbit, Geosynchronous Orbit, Trans-Lunar Injection and to the Earth-Moon L1 point,” the presentation continued.
Based at the Lagrange point L1 – outside of the Earth’s gravity well – a propellant depot would also open up opportunities to international partners for deep space missions – an element of the FY2011 NASA budget.
Launched by a single HLV, the depot could be refuelled using commercial vehicles, reducing the mass required to launch from Earth’s surface on a Lunar or deep space mission via “dry” – and potentially reusable – landers.
“A propellant depot at the Earth-Moon L1 point would significantly improve lunar and deep space exploration mission operations by providing an infrastructure capability for deep space transportation and by opening up participation to international partners and commercial vehicles. This propellant depot outside of Earth gravity well would act as a refuelling station for spacecraft on the way to the lunar surface or Mars.
“A single launch of an HLV with an Earth Departure Stage can deliver a propellant depot to Earth-Moon L1 point. After that launch, any combination of HLV and existing launch systems could then be used to economically transport propellants to this depot creating an in space market place for commercial refuelling of a variety of deep space missions.
“One of the advantages of a propellant depot is that lunar and Mars landers can now be launched dry. This alleviates the lander structure of needing to carry 70 percent of the total wet mass during the launch phase and transfers propellant delivery mass over to commercial launch sector.
“An HLV with an EDS can deliver an Orion CEV and a dry lunar lander to L1. The dry lunar lander would be loaded with 25 mT of propellants at the depot to complete the lunar phase of it’s mission. Dry launch of the LSAM element to L1 dramatically reduces the spacecraft weight constraints, permitting more flexible and robust operational capabilities to be designed into the lander.
“The lunar lander can also be designed for reusability, docking with the propellant depot for refuelling multiple times. Such a reusable lunar lander offers significant advantages in cost per mission, mission operations schedule flexibility and flight risk, along with reducing the requirements for total launch mass. The L1 staging point can be also be used effectively for simplifying a variety of other deep space missions, including robotic and human visits to asteroid or the moons of Mars.”
Propellant depots are also a key element of the United Launch Alliance (ULA) exploration “master plan”, which focuses on utilizing the Centaur derived Upper Stage, refuelled at a L-1 based depot. That same approach is shown via the SD HLV’s “Open Architecture – NASA Crew Elements and Commercial Propellant” example, showing “L1 Enables Anywhere, Anytime, Global Access”.
SD HLV – NEO Mission:
Examining the use of a HLV for a Near Earth Object (NEO) mission to an asteroid, the assessment claims the development schedule for the vehicle would allow for a near-term demonstration of the technologies required ahead of a 304 day mission to Near Earth Object 2001 GP2 – which would require a departure from Earth in 2019.
“The HLV development schedule and payload capabilities could support a near-term human NEO demonstration that would test operational scenarios and technologies needed for later NEO and Mars missions. Figure 5-17 shows the mission scenario for a two-launch Block II HLV option for a mission to Near Earth Object 2001 GP2 requiring an Earth departure in 12/9/2019 and a total mission duration of 304 days.”
Both the side mount and in-line Block II HLV options are classed as viable for such missions – given both have sufficient payload capabilities to support this relatively low delta-V mission. A modified version of ESA’s ATV (Automated Transfer Vehicle) would provide the basis of a Habitat/Propulsion Module, increasing the international flavor to the mission.
“The first launch places an Orion spacecraft with Service Module and a Habitat/Propulsion Module in LEO. The Service Module has larger capacity propellant tanks than a lunar version to conduct the Trans-NEO injection, NEO Arrival, and NEO Departure burns,” noted the presentation, outlining the mission.
“The Habitat/Propulsion Module is a derivative of the ESA Automated Transfer Vehicle (ATV). With the Habitat element the usable volume of the Orion spacecraft is doubled to provide for accommodations for the crew of 4. For this demonstration mission crew size can be adjusted depending on the amount of consumables that can be accommodated.
“The ATV Propulsion Module has been modified to replace a portion of the ATV propellants with water and other required human consumables. The remaining propulsion capabilities are used for the smaller midcourse correction burns, attitude adjustments, plus station keeping and fly arounds while at the NEO. The Habitat element is also radiation hardened for long-term space exposure.
“For an early mission capability, closed loop support systems are minimized, but small prototype systems may be carried for test purposes.”
Again the dual HLV – one manned, one unmanned – architecture is utilized, with the second launch of the EDS launching after the crew, to minimize boiloff effects on this large cyro stage.
“After rendezvous and docking in LEO with the Orion/SM/Habitat, the EDS stage places the combined stack into a High Earth Orbit (HEO) 800 km x 120,500 km. As the stack approaches apogee, the EDS is jettisoned and at the appropriate time the Orion Service Module performs the Trans-NEO Injection burn (0.395 km/sec).
“The NEO arrival burn of 2.073 km/sec takes place 130 days after launch. The primary exploration mission at the NEO lasts 14 days. The NEO departure burn requirement is only 0.170 km/sec. The Earth return phase lasts 160 days. As the spacecraft approaches Earth, the Habitat is jettisoned followed by the Orion Service Module. The Orion then makes a supercircular velocity entry and an ocean landing.”
A second – shorter duration – mission option is also referenced, which aligns with NASA’s own Flexible Path evaluations, based on a mission to the NEO 1999AO10 in 2025-2026.
Also see NASASpaceflight.com’s Flexible Path Review:
Part 1: Battle of the Heavy Lift Launchers – Monster 200mt vehicle noted
Part 2: Manned mission to construct huge GEO and deep space telescopes proposed
Part 3: NASA Flexible Path Evaluation of 2025 human mission to visit an asteriod
Part 4: Taking Aim on Phobos – NASA outline Flexible Path precursor to Man on Mars
“The mission scenario (for) Near Earth Object 1999AO10 (would) launch on January 2, 2026. This mission lasts only 155 days (vs. 304 days for the previous NEO mission), but the total delta-V requirement is higher (4.120 km/sec vs. 2.638 km/sec). This requires the use of the Block III HLV that has a payload capability of ~ 110 mt,” added the presentation, with the slide taken from the Flexible Path presentation.
“For this mission, an inflatable design Habitat described in the NEO DRM (Design Reference Mission) discussion could be used. Closed loop systems, tested in the earlier demonstration mission, would be incorporated. Because of the higher delta-V requirements, the launch scenario is different.
“The first launch carries the EDS stage and Habitat to orbit first. The higher propellant load Orion/SM is then placed in LEO on the second launch. Thereafter, the mission sequence is the same as for the demonstration mission, but with differing delta-V and duration requirements.”
Other SD HLV applications are also referenced in the presentation, such as a prototype Solar Power Satellite (SPS) mission as early as 2016 – launching a 30-mt SPS demonstrator to “underscore” the abilities of the HLV. Such options will be outlined in future articles.