Crew Exploration Vehicle

1
Crew Exploration Vehicle

• Jana Schwartz
• jana_at_draper.com

2
CEV Lunar Direct Return (Arch 1) ConOps
CEV Launch on human-rated LV. Automated launch
abort performed by dual-use IPU (no tower needed).
3
CEV Lunar Direct Return (Arch 1) ConOps
Automated Rendezvous Dock with pre-deployed
stack in LEO. CEV is active member in ARD. If
dock is not achieved in 48h, perform burn,
jettison IPU and return to Earth.
4
CEV Lunar Direct Return (Arch 1) ConOps
Transfer to Moon. CEV provides habitation and
serves as control center for all stack
operations. Jettison EDS after TLI and LOI
burns.
5
CEV Lunar Direct Return (Arch 1) ConOps
Descent to lunar surface. CEV provides automated
or astronaut-in-the-loop control interface for
landing.
6
CEV Lunar Direct Return (Arch 1) ConOps
Lunar surface operations. CEV serves as surface
habitat for short lunar missions (baseline 4
crew x 5 days). One full-crew EVA per day.
7
CEV Lunar Direct Return (Arch 1) ConOps
Ascent from surface and direct return to
Earth. CEV provides habitation and serves as
control center for all TEI operations. Direct
return architecture provides mass-efficient
anytime-abort capability.
8
CEV Lunar Direct Return (Arch 1) ConOps
Midcourse correction. MCC performed by
IPU. Jettison IPU max 12h prior to reentry.
9
CEV Lunar Direct Return (Arch 1) ConOps
Earth entry descent. CEV limits g-loading to
acceptable ranges for non-conditioned
crew. Ablative heat shield mitigates thermal
loads.
10
CEV Lunar Direct Return (Arch 1) ConOps
Terminal landing recovery. CEV alights on land
to simplify nominal recovery process, but can
support contingency water landing.
11
CEV Reference Design for Mars Moon
docking port not required for Lunar direct
return architecture supports Mars missions
12
Integrated Propulsion Unit (IPU) Reference Design
13
CEV/IPU Subsystems
14
Reference CEV/IPU Mass Breakdown
15
CEV Capability Frontiers
CEV push mass is the mass of the CEV/IPU after
ARD with the stack
16
CEV as Short Mission Lunar Surface Hab

• Surface-capable CEV presents more challenges than
  orbiting CEV, but allows for more efficient
  Exploration System of Systems
• Additional requirements levied on surface-CEV
• Support crew during M-landing / launch
• Crew will recline in couches, as required for
  Earth reentry and Mars landing
• Display landing system data for
• astronaut-in-the-loop control
• Can use same display screens as in Earth
  launch/entry ops
• Possibility for synthetic vision techniques with
  manual sighting backup
• Surface access (nominal ingress/egress through
  side door contingency through top hatch)
• Habitation in gravity field
• Dust mitigation within crew cabin
• Discretionary capabilities for surface-CEV
• Airlock
• Minimal air/tank mass does not justify CEV
  airlock, but need to further consider other
  subsystems (e.g., avionics)
• Requirements levied on other modules
• Provide access to surface (e.g., ladder,
  counterweight-assisted rappel system)
• House sensors, actuators, structure for safe
  landing
• Mitigate surface environment, especially thermal

17
CEV Habitable Volume

• Design Space
• What amount of habitable volume is required for a
  crew of 4 for 14 days (maximal Moon) or a crew of
  6 for 5 days (contingency Mars)?
• What restrictions are placed on the size/shape of
  the CEV by the launch vehicle?
• What is the best moldline to support packing of
  the crew and crew systems?
• What crew systems will be required and how will
  they be packaged?

Apollo CM

• Drivers
• NASA
• Satisfy human habitation requirements
• SoS
• Fit on existing launch vehicle
• Support stack configurations
• Support direct return
• Mars-Back flexiblity, extensiblity
• CEV
• Minimize mass
• Package all required systems in allotted volume

• Findings
• Taking into consideration factors modifying crew
  tolerances and activity based models, the
  required habitable volume can be reduced from the
  initial assessment of 17m3
• A capsule design is sufficient to meet the
  pressurized/habitable volume requirements
• Recommendations
• Pressurized volume 27 m3 (purple region)
• Habitable volume 12.8 m3
• Use partial body board for support of
  incapacitated crew member

18
Current Volume Determination Methods
NASA-STD-3000

• Habitable volume requirements derived from
  current methods
• Crew of 4 for 14 days
• 4.3m3 per crew member
• 17m3 total CEV habitable volume

Fraser, 1966 NASA CR-511

• Factors Modifying Tolerance
• Motivation and experience of the subjects
• Ambient environment
• Activities and tasks during confinement
• Knowledge of duration
• Exercise and physical fitness of subjects
• Number of confinees

19
Activities Based Model

• Consider activities to be performed by crew
• Command and control
• Mission accomplishment
• Science
• Exercise/recreational
• Medical/health maintenance
• Incapacitated Crew Member
• Food preparation/eating
• Group gatherings/entertainment
• Sleeping/privacy
• Clothing changes
• Personal hygiene
• Assume 95th percentile male

20
Multi-ConOps Operational Analysis
27m3
12.8m3
Mars, 5 crew
Short Lunar, 4 crew
EVA, 5 crew
21
CEV Flight Critical Operations
to LEO
from LEO / Moon / Mars
EDL Config, TPS, Aerostability
Recovery Sites
late launch abort
reentry
Tower vs. Dual-Use IPU 10g / 3s
early launch abort
Precision Landing
descent
Terminal Landing
landing
22
CEV Abort Option Space

• CEV w/o crew escape 12.7 tonne (28,000 lbm)
• CEV DeltaV 305 m/s (1000 ft/s)
• Abort Average Acceleration 8g
• Trade options considered
• Ejection Seats
• Dual use High Thrust-to-weight engine system
• Solids, dual use
• Solids, independent
• Dual use liquid with low pressure thrusters
• Independent Liquid Abort system
• When considering mass only, Ejection seats and
  Dual use, high T/W system provide best solution

Source Aerojet
23
CEV Pad to Max-q Abort FootprintsAtlas V
Trajectory
10 deg/s Maximum Attitude Rate

• Constraints on the pitch rate can dramatically
  reduce abort footprints

24
Recovery Site Definition

• Extend recovery ship operating range by using
  rescue helicopter
• Maximum allowed crew wait time (0, 1 2 hours)
• Loiter option converts 2 hr wait/range to 1 hr
  out / 1 hr wait (with identical range)

Site 3
Site 2

• Apollo pararescue requirement was 4 hours
  actual 19-45 min (avg. 34 min)
• Helicopter S-61 SeaKing, 265 kph cruising speed,
  8.5 hr max duration

2 hr range/wait
2 hr range/wait
1 helicopter, 5 hr total fuel
Loiter option
1 hr out
1 hr out
1 hr range/wait
1 hr range/wait
1 hr wait
1 hr wait
2 hr range
2 hr range
1 helicopter, 3 hr total fuel
2 helicopters, 7 hr total fuel
25
Entry Vehicle Configuration

• Design Space
• Entry Velocity Nominal LEO (7 km/s), Nominal
  Lunar (11 km/s), Nominal Mars (11.5-12.5 km/s),
  Off-nominal180-day anytime Mars (14 km/s)
• L/D (Mold Line)
• Ballistic L/D 0
• Low-Lift L/D 0.1
• Conic L/D 0.3
• Bi-Conic L/D 0.6
• Lifting Body L/D 1.4
• Ballistic coefficient (CB) 200-600 kg/m2
• Reentry g-Limit 2, 3.5, 5 8 gs

• Findings
• Lifting body provides significantly more corridor
  and trajectory reach, and reduced g-loads at
  entry but at expense of extreme thermal heat
  rate and heat loads
• Mars Off-nominal drives configuration
• Moderate L/D (0.3) satisfies entry corridor and
  5g reentry
• Peak heat rate 1600 W/cm2 (L/D 0.3)
• Max heat load 140 kJ/cm2 (L/D 0.3)
• Recommendations
• Non-lifting body vehicles
• L/D 0.3 to 0.6

• Drivers
• Deceleration
• g-limits for human crew are primary driver for
  entry configuration selection
• L/D, Entry Velocity Ballistic Coefficient
  determine vehicle ability to limit deceleration
• Heating
• L/D, Entry Velocity Ballistic Coefficient
  determine vehicle heat rate and heat load
• Heat rate determines TPS selection (ablative or
  reusable)
• Heat load determines TPS thickness

26
Corridor and L/D Required

• g-limit Considerations
• Mars Off-nominal return requires L/D 0.25 for
  5g limit
• LEO Return L/D 0.2
• Mars Return L/D 0.15
• Lunar Return L/D 0.1
• Minimum required L/D for LEO greater than both
  Lunar and Mars return
• Decreasing dynamic pressure increases required
  L/D to maintain g-limit
• Significantly higher L/D required for extreme low
  g-limits (e.g., 2 gs)
• Increases significantly with entry velocity
• Entry Corridor Considerations
• L/D VEntry are primary drivers for corridor
  width, not ballistic coefficient (CB )
• Mars Off-nominal return requires L/D 0.3 for
  adequate corridor

27
Thermal Environment

• Heating Drivers Vehicle Configuration
• Vehicle nose radius Entry heating rates
  increase with decreasing radius (lifting body ½
  conic)
• Surface area Large area increases radiative
  heating
• Lifting body 3.3x conic
• Dominant effect for VEntry gt 12.5 km/s
• Lifting body heat rate and heat loads are extreme
  as compared to other moldlines considered
• Can be biased towards shallower trajectories
  within corridor to reduce peak heat rates but at
  cost of higher heat loads
• Heating Drivers Entry Velocity
• Non-lifting body heat rates and heat loads are
  essentially identical at nominal Lunar / Mars
  entry velocities
• Heat rate and heat load increase dramatically for
  Mars Off-nominal
• CEV Thermal Environment (L/D 0.3 to 0.6 for
  Mars Off-nominal return)
• Peak heat rate 1200 to 1600 W/cm2
• Max heat load 110 to 140 kJ/cm2

Heating plots show nominal 5g limit
trajectories, flown at center of corridor and
include both radiative and convective heating
28
Trajectory Reach

• Reach is capability, not landed accuracy
• Range definitions
• Downrange is measured in the entry plane from
  entry interface
• Crossrange is measured as distance perpendicular
  to entry plane
• Reach is achieved by applying lift in the
  direction that gives maximum range (downrange or
  crossrange) while staying within g-load
  constraints
• Reach is required when landing site cannot be
  placed within ballistic footprint
• For example, EDL from LEO may not have landing
  site in the descent plane
• Alternate sites may be required when primary is
  unavailable (e.g. weather) after committing to
  entry trajectory from Lunar transfer
• Can be extended with skip trajectories (not
  considered in this analysis)
• Increasing lift (L/D) significantly increases
  downrange and crossrange capabilities
• Downrange footprint Lifting body is 5x conic
• Crossrange footprint Lifting body is 20x conic

29
Thermal Protection System
TPS Options Ablative vs. Reusable

• Drivers
• Heat rate determines TPS material properties
• Peak heat rate lt 1,600 W/cm2 (Mars
  Off-nominal, L/D 0.3, non-lifting body)
• Heat load determines TPS thickness (mass)
• Max heat load lt 140 kJ/cm2 (Mars
  Off-nominal, L/D 0.6, non-lifting body)
• Nose radius and vehicle size also drive
  convective and radiative heating

Ablative

• Findings
• Reusable TPS not viable option at lunar return
  speeds (heat rates gtgt 35 W/cm2)
• Mars off-nominal return (14 km/s) drives TPS
  sizing
• No human-rated ablative available requires
  extensive testing

Reusable

• Design Space
• Conservative TPS sizing analysis performed
• Sizing completed for uniform thickness
• Sized for stagnation-point (most massive option)
• Compared Carbon Phenolic with PICA to bound
  sizing
• Mass estimates are forebody TPS (wrt flow field)
  do not include structure or aftbody TPS

• Recommendations
• Human-rate low/mid density phenolic for peak heat
  rates up to 1,600 W/cm2
• Investigate mass vs cost gains for non-uniform
  thickness and /or hybrid insulator layers
• Material thickness could be optimized by Phase
  (Lunar vs Mars return)

30
TPS Options for Entry
From Moon
From Mars
From LEO
Ablative
Reusable
Courtesy Bernie Laub, NASA Ames Research Center
Fully capable Not capable
Potentially capable, but not demonstrated Potentia
lly capable, but large TPS mass penalty
X

• No single TPS design exists for all Lunar/Mars
  applications
• Recommend program to optimize TPS material(s)
  design (e.g. hybrid layers)

31
Terminal Landing Analysis
Airbags
Crushables

• Findings
• All systems appear feasible
• Deceleration within Brinkley model
• Terminal System mass ranges from 60 to 100 kg
• Crushables are passive, but require most volume
  (crush area by stroke length)
• Airbags use less space, but require dependable
  inflation / deflation mechanisms
• Solid propulsion is heaviest (mounting under heat
  shield would require jettison of heat shield)
• Terminal descent rates less than 6 m/s increase
  parachute mass significantly
• Scaled chute mass 365 kg for 5.4 tonne CEV at
  6 m/s
• Recommendations
• Airbags selected for packaging efficiency
• Addition of floats to keep vehicle upright in
  water
• Assess passive backup systems
• TPS insulation that doubles as crushables
• Couches with custom fitted liners with energy
  absorbers

Crushable behind heatshield
Sandwiched Airbags
Propulsive
Airbags behind heatshield

• Drivers / Design Space
• Propulsive soft landing
• Good control of touchdown velocity
• Solids are lighter weight, but provide less
  control
• Liquids are throttleable, but significant mass
  penalty
• Honeycomb/foam crushables
• Simple, low mass system doubles as TPS
  insulation
• For accelerations below 5gs lower crush strength
  materials or stabilization will be required
• Accelerations below 5gs require very long crush
  strokes
• Airbags
• Good for secondary impact (if the vehicle
  bounces)
• Impact loading is tailored by airbag geometry
• Could place airbags between heatshield and
  vehicle
• Provides known surface for airbags to impact

32
Parachute Descent Mass Trade

• Parachute mass scaled from Apollo based on area
  (3 parachutes)
• Total Apollo main chute mass 203 kg
• Apollo main chute diam. 26.8 m
• System mass sizing is impacted by landing
  velocity
• Design knee at 6 m/s
• Descent rates less than 6 m/s increase parachute
  mass and thus total mass significantly
• Parachute mass sizing will be designed for
  nominal descent rate at 6 m/s
• Terminal landing system mass sized for 1 chute
  failure at 7.2 m/s (previous chart)
• For 5.4 tonne CEV,
• Scaled chute mass 365 kg
• Airbag landing system mass 85 kg
• Additional performance can be gained by trading
  increased descent rate with airbag failure
  tolerance

Nominal (6 m/s)
Total
Apollo (8.2 m/s)
Chute
Airbags
Total
5.4 tonne
Chute
Airbags
33
Draper/MIT CEV