why are we doing this again?
DESCRIPTION
1 - Introduction 2 - Propulsion & ∆V 3 - Attitude Control & instruments 4 - Orbits & Orbit Determination 5 - Launch Vehicles Cost & scale observations Piggyback vs. dedicated Mission $ = 3xLaunch $ The end is near? AeroAstro SPORT. 6 - Power & Mechanisms 7 - Radio & Comms - PowerPoint PPT PresentationTRANSCRIPT
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Why are we doing this again?
• 1 - Introduction• 2 - Propulsion & ∆V• 3 - Attitude Control
& instruments
• 4 - Orbits & Orbit Determination
• 5 - Launch Vehicles– Cost & scale
observations– Piggyback vs. dedicated– Mission $ = 3xLaunch $– The end is near?– AeroAstro SPORT
• 6 - Power & Mechanisms • 7 - Radio & Comms• 8 - Thermal /
Mechanical Design. FEA
• 9 - Reliability• 10 - Digital & Software• 11 - Project
Management Cost / Schedule
• 12 - Getting Designs Done
• 13 - Design Presentations
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Orbiting down memory lane...• Kepler & Conics (Mostly
Elipses)• Period, Velocity, Radius, Escape• Orbit descriptions: (6)ephemerides• Orbit transfers: Hohmann• Gravity assist: M motion Matters• Harmonic, frozen, synchonous
orbits• Oblates, Prolates, J-2 and sun synch• Lagrange Points (stable & un)• GPS: 4 equations, 4 unknowns• Speaking of Oribits:
– Nutation; Precession; Nodes; Line of nodes; Semi-major axis, “The Paramter, P”, Right Ascension, Argument of Perigee, True Anomaly, Vernal Equinox, Inclination, Azimuth/Elevation/Declination, Geoid, Periapsis / Apoapsis, Julian v. Gregorian
• Sidereal day & Geosynch
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Return on Investment
-25000
-20000
-15000
-10000
-5000
0
5000
10000
0 5 10 15 20 25 30 35
Month
Month
Revenue - Investment
(revenue - investment)
Investment Value (with i)
But first, a word from our sponsor: $$$
A large number of small monthly payouts ------
…adds up to a lot of negative equity ------
…and even more with foregone interest included ------
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You Are Here
Design Roadmap
DefineMission
ConceptSolutions &Tradeoffs
ConceptualDesign
Requirements Analysis
OrbitPropulsion
/ ∆VComms
AttitudeDetermine & Control
LaunchGroundStation
Thermal /Structure
Deployables
InfoProcessing
Top Level Design
Iterate Subsystems
Suppliers / Budgets
PartsSpecs
Mass
Power
$
∆V
Link BitsMaterialsFab
Detailed DesignFinal Performance
Specs & Cost
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For next time• Requirements Doc
– Mission Requirements
– System Definition– Begin Tech
Requirements
• Launch Strategy– Primary LV and cost– The last mile
problem
• Reading– Requirements Doc
Sample– Power:
• SMAD 11.4• TLOM 14
– Mechanisms:• SMAD 11.6 (11.6.8 too)• TLOM ?
– Fill in re ACS: TLOM:• Chapt. 6 (magnets)• Chapt. 11 (ACS)
• Thinking– What can you
build?– What can you test?
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Req. # Requirements Crit. Valid. Appr. Source/Comments1. Mission1.1 Piggyback payload insertion orbit shall be dependent on primary payload orbit 3 Review1.2 Primary payload orbit shall not be affected by piggyback orbit
3Demonstration,
Analysis, Review
1.3 Piggyback payload shall be given 0.6 - 1.2 m/s separation velocity relative to the launch vehicle.Velocity vector shall be along payload longitudinal axis
2 Review
Shuttle Hitchhiker is capable of 0.3-1.2 m/s while the ASAP 5 has the maximum velocity (3 m/s) capability. This and CG data will be required for Tip-off and Collision Avoidance Analysis
1.4 Piggyback payload shall be designed for operation without special orientation or spin required atseparation from launch vehicle 2 Review
Some launch vehicles may be able to provide special attitudes and spin at separation, but commonality defines this requirement
1.5 A representative Dummy Payload shall be provided to launch vehicle provider at the beginning ofthe launch campaign
3 Delivery
1.5.1 The dummy payload shall be flight worthy 2 Demonstration1.5.2 The dummy payload shall be representative of the actual piggyback payload in terms of
mechanical interface2 Demonstration
1.5.3 The dummy payload mass shall be within .5 kg of the piggyback payload mass 2 Demonstration1.5.4 The dummy payload shall have the same CG as the actual payload 2 Demonstration1.5.5 The dummy payload shall take the actual payload's place if it is not ready for integration and
launch 3 Delivery
This allows the launch vehicle provider the maximum leeway in launching without the active piggyback payload in case of schedule delays
1.6 Only one person per piggyback payload will be allowed in launch center during launch1 Review
This is only required by Arianespace for ASAP 5. Other launch vehicle providers may be more lenient
1.7 There shall be no standard access to piggyback payload after encapsulation 2 Review1.8 Piggyback payload shall have lifting points for handling and movement of satellite during ground
operations, transport, and encapsulation1 Inspection
2. Schedule2.1 Launch schedule shall be driven by the primary payload and ONLY the primary payload 32.2 Nominal mission start shall be 40 months before launch (T- 40 months) 2 Review On some missions, notably STS, there
might be some leeway2.3 Application for piggyback use on the launchers shall be a minimum of 40 months before launch (T -
40 months)2 Review
2.4 Interface Control Document (ICD) shall be completed for review by launch vehicle provider aminimum of 25 months before launch (T - 25 months)
1 Review See Documentation section for documents that are also due at the same time
2.5 All piggyback payload testing shall be completed as required in the Validation/Testing section aminimum of 7 months before launch (T - 7 months)
2 Review
2.6 Piggyback payload shall be ready for delivery to launcher integration site for integration with launchsystem 6 months before launch (T - 6 months)
3 Review
2.7 Piggyback payload shall be fully tested, fueled and mission ready for integration with primarypayload-launcher combination a minimum of 90 days (T - 90 days) before launch. 3 Inspection
This requirement is driven by STS Hitchhiker. There may be some leeway for piggyback payload provider
2.8 If required, a Structural Integrity Verification Report shall be ready for review by launch provider atleast 13 months before scheduled launch (T - 13 months)
2 Review
This is required of piggyback payloads on Shuttle Hitchhiker. There is some leeway on Shuttle launch schedules, so delays may be negotiated with NASA. This requirement may be ignored on other launchers
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2.0 System Definition2.1 Mission Description2.2 Interface Design
2.2.1 SV-LV Interface2.2.2 SC-Experiments Interface2.2.3 Satellite Operations Center (SOC) Interface
3.0 Requirements3.1 Performance and Mission Requirements3.2 Design and Construction
3.2.1 Structure and Mechanisms3.2.2 Mass Properties3.2.3 Reliability3.2.4 Environmental Conditions
3.2.4.1 Design Load Factors3.2.4.2 SV Frequency Requirements
3.2.5 Electromagnetic Compatibility3.2.6 Contamination Control3.2.7 Telemetry, Tracking, and Commanding
(TT&C) Subsystem3.2.7.1 Frequency Allocation3.2.7.2 Commanding3.2.7.3 Tracking and Ephemeris3.2.7.4 Telemetry3.2.7.5 Contact Availability3.2.7.6 Link Margin and Data Quality
3.2.7.7 Encryption
(Some) STP-Sat Requirements
NB: this is an excerpt of the TOC - the entire doc is (or will be) on the class FTP site
Requirements & Sys Definition go together
Highly structured outline form is clearest and industry standard
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Launch Vehicles
> Review Propulsion and ∆V requirement
> Performance and staging
> Practical Considerations
>Cost & scale observations
>Piggyback vs. dedicated
>Mission $ = 3xLaunch $
>The end is near?
> AeroAstro SPORT
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∆V = gIspln(R)
∆V = ∑i {Vi∆mpi/(M(p))} => V∫{dm/M} (from M=Mo to M=Mbo)
= Vln(M/Mo) = gIsp ln(mo/mo-mp) = gIspln(mo/mbo) = gIspln(R)
Where gIsp includes pressure effects; R is the mass ratio: mass(start)/
mass(burnout)
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∆V = gIspln(R):
Staring at logarithmic
reality 3000
2000
1000
0
∆V
meters per sec
Propellant mass (kg)0 10 20 30 40 50
∆V Performance Samples: dry mass 50
kg
Isp 300 seconds
Isp 60 seconds
Staging is an
answer...
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Single vs. Two
Stage Assumptions: • R = M(i)/M(f) = 10
• ∆V required: 10 km/s
• Payload = 100 kg • Payload =10% MfSSTO: 100 kg payload
∆V = gIspln(R):
Isp = 420 (H2 / O2)
Launch mass: 12,500 kg
Structure = 1000 kg
=> R = 12.5
Stage payload Mass Fraction: 0.8%
TwoSTO: S-1 ∆V(s)=5000m/s (2 stages, equal ∆V)
S-2 mass: 505 kg
S-2 structure: 150 kg
S-2 PMF: 20%
TwoSTO: S-2 ∆V(s)=5000m/s
S-1 mass: 2595 kg
S-1 structure: 770 kg
S-2 PayMF: 20%
TwoSTO: ∑ ∆V =10000m/s
Total Mass: 3100 kg
Total PayMF: 3.2%
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Costs of Orbital Insertion
• Naïve Observations:– Bigger rockets are
cheaper, regardless of who builds them
– ‘50s technology Scout costs @ same as ‘90s technology Pegasus
– Bringing things back from orbit and/or crewed vehicles cost more
– Marginal cost to fly a 10 kg payload is $50k.
Payload / kg
20,000
10,000
5,000
15,000
Ariane VProton
X
Delta /LLV
X
Pegasus / Scout
X
0
104103102
25,000Shuttle
(est.)
X
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Launch Costs vs. Mission Costs
• Rationale
– Add features to achieve cost parity
– Add standards to achieve cost parity
• MIL-Spec parts, testing...
– Increased launch cost motivates:• Risk Avoidance
– MIL and S-Class Parts– Redundancy– More quality control
» Staff + procedures• Higher value missions
– Multiple payloads– More capable spacecraft
» Pointing, power, data rate
– Parity between launch sponsor and spacecraft sponsor
– Ops cost = Satellite Cost = Launch Cost
• Numbers
– Satellite Cost = Launch Cost
– Scout / Pegasus Payloads• ALEXIS + REX: $24M• HETE / SAC-B: $25M• Microsats: $6M• REX / TEX: $6M• Stacksat $6M• 8 x Orbcomm $24M• MSTI-2 $14M
– Ariane ASAP class payloads• Amsat Oscar $200k (typ.)• Oscar 13 $200k• 4 x Microsats $200k• Astrid (Kosmos)$1M
– Ariane / Long March Interstage
• Freja $4M
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AMSATs piggybacked on Ariane
Oscar 13 (L) cantilevered by a marmon clamp to the payload
adapter ring and a UoSat (below) being prepared for mounting on
ASAP ring
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New Options to Orbit
• Candidates– Aircraft: carry, balloon,
tow
– SSTO: autogyro, Shuttle-like, DC-X, Suborbital
– Sea Launch
– “Cheap” Russian rockets
– Reusable rockets
– “Cheap” US, Indian, Spanish, Brazilian, Chinese or Italian rockets
• Perspectives– Jet Aircraft / Ford
(Taurus) costs over last 40+ years
– Pegasus v. Scout
– AF EELV cost goals (marginal savings)
– Labor cost distortions
– Commercial Competition: Ariane v. Long March v. Proton v. Delta
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Space Transportation’s Future
(15 year outlook)
• Per kg cost may slowly decrease (5% or 10%) - mainly due to competition from new entrants
• Reliability is key, not $/kg• Payload mass (for same
performance) decreasing by 10x per decade– (though large payloads will not
shrink)
• Space Tourism, but suborbital (excepting special cases)
• More use & availability of piggybacks and multiple payload launches
• upper stages replaced by on-board electric propulsion
• Wildcards: siting and environmental issues
• Low cost components ≠ low cost rockets:
hardware vs. reliability $
Hint: Nobody lives at the north pole, and launches won’t cost $10/kg
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The Next Generation of Microspace
Small Payload
ORbit Transfer
TM
AeroAstro Proprietary
TM
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What is SPORT?
Small PayloadORbit Transfer
SPORTMicrosatellite Going to Custom Orbit
Microsatellite Going to GTO(No SPORT)
Enc
ount
er \
SA
IC
Arianespace
Ariane 5Heavy Launcher
Upper StagePropulsion
Upper StagePropulsion
TM
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SPORT GTO to LEO Transfer
1
2
3
4
5
6
123456
Launch into GTOPerigee lowering burnAerobraking drag near perigeeApogee reduction with each passPerigee raising burnFinal circular orbit
SPORT™ Microsatellite
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Aerobraking• Highly efficient orbit transfer (over 2 km/s ∆V)
• Rarified atmosphere altitude - minimal heating
• Large deployable increases profile area ( 50)
• ~ 200 passes to lower apogee 35,000 km
• Nominal 30 day mission
1
2
3
4
5
6
1 Launch into GTO2 Perigee lowering burn3 Aerobraking drag near perigee4 Apogee reduction with each pass5 Perigee raising burn6 Final circular orbit
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SPORT Releases Microsatellite
Dispose SPORT™
Release Microsatellite in Custom Orbit
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0
500
1000
1500
2000
2500
3000
3500
4000
4500
5000
0 5 10 15 20 25 30Maximum Inclination Change, deg
48 kg 41 kg 26 kg 14 kg
Initial GTO Orbit620 x 35,883 km altitude
7 deg inclinationAerobraking for Apogee Reduction
Payload Mass
Aerobraking PerformanceUtilizing the aerobraking and propulsion features of SPORT, a wide range of missions is possible.
Note: Assumes total initial mass of 100 kg.
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SPORT ™ performs a variety of orbit transfer maneuvers
GTOTo
GEO
LEO to MEO
Sun Centered
GTO to LEO
Molniyato SSO
L4
L1
L2
L5
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Molniya to SSO Transfer• Initial Orbit: Molniya
– 510 km 40,000 km and 62.8 deg
– Launch on Molniya as Secondary
• Final Orbit: – 800 km Sun Synchronous
• SPORT™ Transfer– 900 m/s ∆V Apogee Burn
• 35.8 deg Inclination Change• Lowers Perigee to 150 km
– Aerobraking • Reduces Apogee to 800 km
– 180 m/s ∆V Apogee Burn• Raises Perigee to 800 km
Nominal Payload CapabilityMicro SPORT: 20 kgMini SPORT: 60 kg
Nominal Payload CapabilityMicro SPORT: 20 kgMini SPORT: 60 kg
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LEO to MEO Transfer• Initial Orbit: Polar LEO
– 800 km 800 km and 98.6 deg
• Final Orbit: Polar MEO– 1600 km 1600 km and 98.6 deg
• SPORT™ Transfer – 190 m/s ∆V Perigee Burn
• Raises Apogee to 1600 km
– 190 m/s ∆V Apogee Burn• Raises Perigee to 1600 km
Note: no aerobraking hardware required
1
2
3
4
123
4
Launch into SSOPerigee burnApogee burn
Final circular orbit
Nominal Payload CapabilityMicro SPORT: 50 kgMini SPORT: 150 kg
Nominal Payload CapabilityMicro SPORT: 50 kgMini SPORT: 150 kg
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Direct Transfer PerformanceUtilizing just the propulsion feature of
SPORT, a wide range of missions is still possible.
Note: Assumes total initial mass of 100 kg and aerobraking hardware removed.
0
200
400
600
800
1000
1200
1400
1600
1800
2000
0 1 2 3 4 5 6 7Inclination Change, deg
48 kg 41 kg 26 kg 14 kg
Initial Orbit400 x 400 km
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High Energy Missions• Initial Orbit: GTO
– 620 km 35,883 km and 7.0 deg– Launch on Ariane 5 in ASAP Slot
• Final Orbit Options:– Earth Escape– Lagrange Point– Lunar Transfer– Asteroid Flyby
• SPORT™ Transfer V Burn at Perigee
L4
L1 L2
L5
Nominal Payload CapabilityMicro SPORT: 20 kgMini SPORT: 60 kg
Nominal Payload CapabilityMicro SPORT: 20 kgMini SPORT: 60 kg
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SPORT Systems
Bitsy kernel• Developed for NASA and USAF• Includes core satellite capabilities
- Communications- C&DH- Power regulation- G&C
Batteries• Variety of options
based on flight proven technology
Aerobrake• Provided by proven supplier• AeroAstro patent pending• Modular per mission
Propulsion System• Modular per ∆V required• Simple spin stabilized design
Microsatellite PayloadPayload Interface Ring