Technology Payoffs for Human Space Exploration - PowerPoint PPT Presentation

Loading...

PPT – Technology Payoffs for Human Space Exploration PowerPoint presentation | free to download - id: 6b84b4-YmE0M



Loading


The Adobe Flash plugin is needed to view this content

Get the plugin now

View by Category
About This Presentation
Title:

Technology Payoffs for Human Space Exploration

Description:

Technology Payoffs for Human Space Exploration Presentation for the NASA Technology Roadmaps Workshop Pasadena California March 24, 2011 Gordon Woodcock, Huntsville ... – PowerPoint PPT presentation

Number of Views:21
Avg rating:3.0/5.0
Date added: 21 May 2020
Slides: 25
Provided by: gordonw82
Category:

less

Write a Comment
User Comments (0)
Transcript and Presenter's Notes

Title: Technology Payoffs for Human Space Exploration


1
Technology Payoffs for Human Space Exploration
  • Presentation for the NASA Technology Roadmaps
    Workshop
  • Pasadena California
  • March 24, 2011
  • Gordon Woodcock, Huntsville Alabama,
    grw33_at_comcast.net

2
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.

3
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, Whats 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.

4
The Brute Force (Apollo) Approach to Exploration
  • Implementers of Apollo really didnt 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.

Approach Problems
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
5
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

6
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.

7
Outpost Concept and Landing Configuration
  • 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.

Habitat and Node with Airlocks
Twin Landers with Habitat Node
8
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.

9
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.

10
Whats L1?
?
Moon
  • L1 is a so-called Lagrange point or libration
    point, where gravity of the Earth and Moon and
    the rotation forces due to the Moons rotation
    about the Earth all balance out. Its 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 its a balance point, the propulsion
    required to stay there is very slight, and it is
    a good place for a way station. Its 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.

L1
Earth
11
Transportation Features and Concepts
Crew Lander/Ascent Vehicle
The tug trip is a slow spiral, typically 6
months. The crew trips are about 5 days each way.
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.
  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

12
Representative Manifest
Maintain Outpost
  • 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.

Annual outpost support 9.8 launches including
SEP tug replacement every five uses
13
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.
14
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.

15
Mission Design Challenges of Mars1. 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.
16
Mission Design Challenges of Mars2. 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
    cant 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.

17
Polar Orbit Station Concept
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
  • 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.

18
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.

19
Mars Lander ConceptRe-usable operations fueled
in orbit for landing and on Mars for ascent.
Movable aero surfaces for pitch trim during aero
descent.
LOX-LH2 propulsion modules fore and aft for
balance attitude control by differential
throttling. No gimbals. Engines 115 kN (26
klbf) each, throttling range 31 to
41.
Cargo space
  • 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.

Mars atmosphere is so tenuous that for ascent,
the vehicle can simply fly sideways doesnt
need engines in the tail.
20
Mars Results
  • Analysis of the various Mars mission profiles is
    incomplete (paper isnt 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.

21
The Game-Changing Technologies
Technology Area Requirements 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 gt 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
22
Three More for Mars
Technology Area Requirements 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 Earths. 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.
23
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 isnt 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.

24
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.
About PowerShow.com