Presentation on theme: "Technology Payoffs for Human Space Exploration Presentation for the NASA Technology Roadmaps Workshop Pasadena California March 24, 2011 Gordon Woodcock,"— Presentation transcript:
Technology Payoffs for Human Space Exploration Presentation for the NASA Technology Roadmaps Workshop Pasadena California March 24, 2011 Gordon Woodcock, Huntsville Alabama, email@example.com
Objectives of this Briefing Examine technologies’ importance from the top down instead of by technology area Identify high priority technology areas for NASA. Provide a sense of value in terms of payoffs, risk, technical barriers and chance of success. Specific technologies as “Game Changing Technologies”? I assert this is the only way to shed light on these three questions.
Reasons and Constraints for Human Exploration Why are we doing this? (goals) What do we want to do? (objectives) How best to do it? (means, i.e. systems and architectures) What technologies enable us to meet goals and objectives? As a taxpayer, “What’s in it for me?” – “Our goal is the capacity for people to work and learn and operate and live safely beyond the Earth for extended periods of time, ultimately in ways that are more sustainable and even indefinite.” Introduction to Obama Space Policy, June, 2010 – “… bringing the inner solar system into our economic sphere” … John Marburger, 2006 – Protect the Earth from asteroid/comet impacts The Moon is the most practical, probably only place to develop these capabilties in a real environment. Other places too inaccessible. Constraint: It must be affordable. Sustainability requires costs to be less than benefits. Presently, costs are much too high.
The Brute Force (Apollo) Approach to Exploration Implementers of Apollo really didn’t have a choice. Cost was not an issue. Four approaches were considered # 4 was “least of evils” choice, still brute force; the only “can do” issue could be, and was, solved by a flight program: Gemini. ApproachProblems Assembly in Earth Orbit No support facilities in orbit No experience with rendezvous or docking No basis for estimating a timeline Land Apollo capsule directly on the Moon using storable propellants Required a huge “Nova” rocket; no facility or tooling could handle it Land Apollo directly with cryo propellants in spacecraft No flight experience with liquid hydrogen; no technology for storage Use lunar orbit rendezvous with storable propellants; build a lander module No experience rendezvous or docking; but needed only with small spacecraft
Today is Not The Same Situation As Apollo …. Not “Get there before the Russians” but “Reduce the Deficit and Grow the Economy” Reduce development costs – Common generic solutions – Reduce launch requirements; use more, smaller launches; derivatives of current and commercial launchers Make in-space systems re-usable and more efficient – Take advantage of advanced (electric) propulsion – Use “gas station” concept (propellant depots at destinations, supplied by solar-electric tugs) – Also deliver cargo to gas stations with solar-electric tugs – Use high-thrust propulsion only for human crews
Representative Lunar Project (Try these principles out and see how well they work) Establish a small human outpost (4 – 6 people) that can be continuously staffed, and begin the development of lunar resource exploitation (as well as scientific exploration). Crew exchange cycle on 6-month intervals. Two logistics flights per year. Modified EELVs & competitors: Delta IV H, Atlas V H, Falcon 9 H, ATK-EADS Liberty; uprate to 40 t. to LEO & 6.5 m. fairing Launch rates on the order of one a month, reduce prices significantly below current levels.
Outpost Concept and Landing Configuration Habitat and Node with Airlocks Twin Landers with Habitat & Node The outpost habitat and node are like ISS modules, a bit larger diameter because people need to stand up inside in lunar gravity. Total mass of the two about 25 t., crew 4 to 6. Because of the launch vehicle size limitation (6.5-meter fairing) the landers and the payload accomm- odation are very different in appearance than traditional concepts. Lander internal arrangement appears on a later slide. Landers fly under their own power only in vacuum, permitting the unconventional arrangement.
Mission Architecture Outline Servicing stations in Low Earth Orbit (LEO) and at Lunar L1. Space vehicles sized to be launched intact within launch constraints (No assembly of vehicles in space except berthing or docking.) Solar electric propulsion (SEP) tug solar arrays deployable. – One exception: yoking two landers and a payload together for lunar landing; done at L1 station; bridging structure assembled and installed using pre-fabricated graphite composite struts with quick-lock fittings. Propellants and cargo transported from LEO to L1 by SEP tug. – LEO station used for servicing SEP tug(s) and mating payloads to tug. Crews launched to and returned from L1 in an Apollo-like crew vehicle (Crew entry vehicle, CEV) capable of Earth entry and landing. The Orion now in development is a CEV. – Single direct launch from Earth. Lunar landers single direct launch from Earth, self-powered to L1, refueled at L1 for lunar landing and return, and retained at L1 for refuel and re-use after first use.
Mission Profile Features Electric propulsion trips from LEO to L1 and return are slow spirals, about 6 to 9 months up and 2 to 3 down. – Low fuel consumption (high Isp) comes at the expense of low thrust. Crew trips are high-thrust, about 5 days. (It takes a little longer to go to L1 than to low lunar orbit.) Trips from L1 to the surface and return take a couple of days each way. L1 is an operations base, where crews transfer between CEV and lunar lander, propellants are transferred from transport tankers to storage tanks and from storage tanks to vehicles, and cargos are loaded on or into lunar landers – It has solar electric power, electric propulsion for stationkeeping, propellant storage tanks, docking and berthing ports, and a small emergency crew habitat for 4 to 6 people. Outpost includes a 75-kWe solar electric/regenerable fuel cell power system, as well as exploration and surface operations equipment.
What’s L1? Earth Moon L1 L1 is a so-called Lagrange point or libration point, where gravity of the Earth and Moon and the rotation forces due to the Moon’s rotation about the Earth all balance out. It’s between the Earth and Moon; another, L2 is behind the Moon as seen from the Earth. There are three others far away from the Moon and not of much interest for this mission. Because it’s a balance point, the propulsion required to stay there is very slight, and it is a good place for a way station. It’s always accessible from Earth orbit and any place on the lunar surface is always accessible from L1. It requires somewhat more propulsion performance but if there is a propellant depot there for re-fueling, using L1 as a way station is a net payoff. The picture at the right is what a trajectory from Earth to L1 looks like if the picture is rotating with the motion of the Moon around the Earth so that the Moon appears stationary.
Transportation Features and Concepts 1.Solar electric tug delivers propellant to depot at L1, lander parked at L1, loads propellant. 2.Mission vehicle (CEV) launched to L1, rendezvous with tug & lander 3.Lander refuels at depot and executes surface mission, entire lander returns to L1 4.Crew transfers to CEV and returns to Earth 5.(Not shown) Tug returns to LEO L1 1 4 3 2 Depot The tug trip is a slow spiral, typically 6 months. The crew trips are about 5 days each way. Crew Compartment LOX Tank 1 0f 4 Hydrogen Tank 1 0f 2 Engine 1 0f 4 This mission profile, using electric propulsion and propellant depot at L1, performs crew lunar landing for about half the launch mass of Apollo brute- force approach, and the lander is re- usable. When lunar propellant is put in production, launch mass decreases further. Crew Lander/Ascent Vehicle
Representative Manifest Maintain Outpost Annual outpost support: 9.8 launches including SEP tug replacement every five uses Establishing the L1 station and delivering the outpost to the lunar surface requires 19 launches at 40 t. capacity and 6.5 meter fairing, including effects of inability to always load to 40 t. and large- volume low-density payloads. When lunar propellant production starts, launch requirements are reduced. Lunar propellant production needs equipment and another power system – solar electric lunar day only about 250 kW, same size array as the outpost power system.
Results: Cost Savings Potential Annual Operating Cost Note: Crew launches are lumped in with other launches for the new technology architecture because the same launcher is used for all. For the right-most case, the Orion was assumed re-usable.
Representative Mars Project Establish a Mars orbit station in elliptic polar orbit Establish small human outpost (4 – 6 people) on the surface, intermittently staffed at first; begin far-ranging scientific exploration; develop self-sufficiency through resources exploitation. Increase crew stay times from about a month (first mission) to a Mars synodic period (~ 2.2 years) after several visits. Mission profiles change with increasing stay times – Short stay; long stay; semi-cycler perhaps using electric propulsion Logistics deliveries every Mars synodic period. Launch capability similar to that for the Moon but higher launch rate.
Mission Design Challenges of Mars 1. For short-stay (~30 d.) missions, propulsion requirements vary greatly. “Easy Year” (2032): Visit Mars when Mars is closest to Earth “Hard Year” (2028): Visit Mars when Mars is farthest to Earth Long-stay missions always low energy but must stay at Mars for about 16 months. “Low energy” is still significantly more than going to the Moon.
Mission Design Challenges of Mars 2. Mars missions are long-duration Apollo 11 went to the Moon and back in about a week. Minimum duration for a Mars mission is about 14 months. (Short stay, easy year) Interplanetary transfers are long, 5 to 9 months each way. A long- duration habitat is required; can’t get by with a small capsule as did Apollo. A short-stay mission can park for 30 days in a Mars orbit and do a brief landing, a few days like Apollo. However, this means spending over a year on a space voyage for just a few days actually doing what you went to do. Half or more of a long-stay mission is spent at Mars. If you want to spend this time on the surface, you need 100 to 150 tons of facilities and equipment landed on Mars to support the mission. All in all, a mission to Mars is on the order of ten times as challenging as a mission to the Moon.
Polar Orbit Station Concept Sketch to scale, with Mars’ axis tilted 24.5 degrees to the ecliptic. The parking orbit periapsis is at the pole. 90 degree inclination, no nodal regression. Apsidal advance 0.265 deg/day is nulled, about 1.25 m/s per day. Orbit plane rotated around its major axis to align with arrival and departure vectors, usually near the ecliptic plane. Electric propulsion at apoapse, where the orbit velocity is about 287 m/s. These corrections are readily made by electric propulsion. Mars architectures have always used ad hoc parking orbits. We want a fixed station in Mars orbit Crew safe haven with habitat Operations base with propellant storage and vehicle parking Way station for semi-cycler mission profile
Mission Architecture Summary Deliver Mars orbit station by electric propulsion Deliver propellant to station; modicum of surface infrastructure, SEP tugs plus direct entry landing. Tug returns to L1. First human mission, short stay, easy year or Venus gravity assist. Brief landing 1 – 2 days up to 10 – 15 days depending on status of surface infrastructure. Deliver more infrastructure to surface, including long-duration habitat system with adequate electric power system. Series of one or more long-stay missions. Propellant in production Begin semi-cycler missions – Enable near-continuous human presence on Mars surface, enhances local agriculture potential – May require electric propulsion for difficult years.
Mars Lander Concept Re-usable operations fueled in orbit for landing and on Mars for ascent. Landed cargo can include built-in crew habitat; expendable and re-usable cargo landers Early crew: expendable crew ascent stage and lander Later crew: crew module, re-usable crew lander; this version must refuel on Mars due to delta V and need for thermal protection. Movable aero surfaces for pitch trim during aero descent. Mars’ atmosphere is so tenuous that for ascent, the vehicle can simply fly “sideways”; doesn’t need engines in the tail. LOX-LH2 propulsion modules fore and aft for balance; attitude control by differential throttling. No gimbals. Engines 115 kN (26 klbf) each, throttling range 3:1 to 4:1. Cargo space
Mars Results Analysis of the various Mars mission profiles is incomplete (paper isn’t due until June). Earlier studies without rigorous accounting of launches and vehicle use indicated savings similar to the lunar outpost. Sources of savings: – Two depots; one at L1, one at Mars orbit station. – Solar electric tug delivery of vehicles (except crew trips), cargoes and propellants to L1 and Mars orbit station. – Reduction in launch requirements follows. – All in-space vehicles re-used except Mars lander and CEV; significant because the interplanetary crew vehicle includes a long-duration habitat and interplanetary propulsion system. – The Mars lander can become re-usable when propellant is in production on Mars’ surface. Total program activity level as measured by launch requirements appears to be 2 - 3 times the lunar outpost.
The “Game-Changing” Technologies Technology AreaRequirements and Issues Solar Electric Propulsion Scale-up to 250 kWe (Array size about like ISS) Certify 50-kW class thrusters for flight Multiple LEO-L1 trips for operating cost savings Radiation resistance (well-shielded concentrator cells) Low mass Propellant refrigeration for zero boiloff Hydrogen and oxygen refrigeration machines capable of removing ~ 10 - 20 watts of heat from cryo liquids Multilayer insulation 50 – 100 layers Efficient propellant transfer from one tank to another Operation in zero g Transfer > 95% of the liquid Bladders not practical for cryogenic liquids Slow fluid rotation in tanks for settling, low-pressure pumps & heat exchangers to control phase of flows. Solar arrays with regen- erable fuel cell energy storage for surface power (day-night cycle) Efficient fuel cells and high-pressure electrolyzers Efficient gas tanks for hydrogen and oxygen gas storage Large deployable array(s) ~ 250 kWe enables ~ 75 kWe continuous power on lunar surface
Three More for Mars Technology AreaRequirements and Issues Surface Power At least tens of kilowatts; eventually hundreds or more for propellant and other resources production. Global dust storms reduce daylight sun by as much as 80% for many days; make solar-RFC systems doubtful. Nuclear power would work, but for true sustainability on Mars, means eventually reactors must be produced there. Power beaming from space is a possibility. Mars synchronous orbit is about half the distance of Earth’s. Millimeter-wave RF power should get through the dust storms OK; no rain to stop it. A transmitter/reflector system similar to but bigger than current comsat technology appears feasible. Planetary Landing Maneuverable aerodynamic lander. Attitude control upon landing engine start (supersonic). Cryogenics Tanks need lightweight vacuum jacket; Mars’ atmosphere is condensible at cryogenic temperatures. Pressure very low.
Comments on Launch Processing Technology Helium: Need to reduce use. – Expendable vehicles: Pressurize propellants with propellants; hydrogen over hydrogen; GOX over LOX; nitrogen over kerosene. – Reusable vehicles: Consider not purging tanks on landing. (It isn’t done on airliners.) – Develop tank-exit boost pumps to reduce required in-flight tank pressure, to compensate for greater mass of non-helium pressurant. Launch Processing: Need to prepare for higher launch rates. – Rates predicted, about 12 a year of Delta IV H class for a lunar outpost. Could grow to perhaps twice this for a Mars outpost. – Reduce on-pad time (or more pads) – More automation of checkout processes.
Summary and Conclusions Technologies recommended have significant leverage toward reducing cost of human exploration missions – Eliminate need for large heavy-lift and reduce launch requirements – Most in-space systems re-usable. – A modest number of systems used in different combinations as appropriate enables all near and mid-term exploration missions. – On path to affordability and sustainability; increases likelihood of reaching long-range goals of human operations beyond Earth orbit. All except cryogenic liquids transfer in space and power beaming have some flight experience. These technologies merit near-term funding and fast track to flight demonstration projects, plus supporting mission analysis to focus requirements.