Transcript No Slide Title

Why are we doing this
again?
• 1 - Introduction
• 6 - Power &
Mechanisms
• 2 - Propulsion & ∆V
• 3 - Attitude Control & • 7 - Radio & Comms
• 8 - Thermal /
instruments
Mechanical
• 4 - Orbits & Orbit
Determination
• 5 - Launch Vehicles
•
•
– Cost & scale
observations
– Piggyback vs. dedicated •
– Mission $ = 3xLaunch $
– The end is near?
– AeroAstro
SPORT
•
Enginering
176 #5
Design. FEA
9 - Reliability
10 - Digital &
Software
11 - Project
Management
Cost / Schedule
12 - Getting Designs
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 b
orbits
Oblates, Prolates, J-2 a
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,
Enginering 176 #5
Azimuth/Elevation/Declination, Geoid,
r
v
rp
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 -----Enginering 176 #5
Design Roadmap
You Are
Here
Define
Mission
Concept
Propulsion
/ ∆V
Comms
Solutions &
Tradeoffs
Attitude
Determine
& Control
Launch
Conceptual
Design
Requirements
Ground
Station
Thermal /
Structure
Deployables
Analysis
Info
Processing
Orbit
Top Level Design
Parts
Specs
Materials
Fab
Mass
Suppliers / Budgets
Power
∆V
Link
Bits
Iterate Subsystems
Detailed Design
Enginering 176 #5
$
Final Performance
Specs & Cost
For next time
• Requirements Doc
– Mission
Requirements
– System Definition
– Begin Tech
Requirements
• Launch Strategy
• Reading
– Requirements Doc
Sample
– Power:
• SMAD 11.4
• TLOM 14
– Mechanisms:
• SMAD 11.6 (11.6.8 too)
• TLOM ?
– Fill in re ACS: TLOM:
– Primary LV and
•
cost
•
• Thinking
– The last mile
problem
– What can you
build?
Enginering 176 #5
Chapt. 6 (magnets)
Chapt. 11 (ACS)
Req. #
1.
1.1
1.2
1.3
Piggy back pay load shall be designed f or operation without special orientation or spin required at
separation f rom launch v ehicle
1.5
1.5.1
1.5.2
1.5.3
1.5.4
1.5.5
A representative Dummy Payl oad shall be prov ided to launch v ehicle prov ider at the beginning of
the launch cam paign
The dummy pay load shall be f light worthy
The dummy pay load shall be represent at iv e of the actual piggyback pay load in terms of
mechanical interf ace
The dummy pay load mass shall be within . 5 kg of t he piggyback pay load mass
The dummy pay load shall hav e the sam e CG as the actual pay load
The dummy pay load shall take the actual pay load's place if it is not ready f or integration and
launch
There shall be no standard access t o piggy back pay load af ter encapsulation
Piggy back payload shall hav e lif t ing points f or handling and mov ement of satellite during ground
operations, transport, and encapsulat ion
2.
2.1
2.2
Schedule
Launch schedule shall be driv en by t he primary pay load andONLY t he primary payload
Nominal mission st art shall be40 mont hs bef ore launch (T- 40 months)
2.3
Application f or piggy back use on the launchers shall be a minimumof 40 mont hs bef ore launch (T 40 mont hs)
Interf ace Cont rol Document (I CD) shall be completed f or rev iew by launch v ehicle prov ider a
minimum of25 months bef ore launch (T - 25 mont hs)
All piggy back pay load testing shall be completed as required in the Validat ion/ Testing section a
minimum of7 months bef ore launch (T - 7 months)
Piggy back pay load shall be ready f or deliv ery to launcher integration site f or integration with launch
syst em6 mont hs bef ore launch (T - 6 months)
Piggy back pay load shall be f ully tested, f ueled and mission ready f or integration with primary
pay load-launcher combination a minimum of90 days (T - 90 day s) bef ore launch.
2.4
2.5
2.6
2.7
2.8
Val id. Appr .
3
Review
3
Demonstrat ion,
Analy sis, Rev iew
2
Review
2
Review
3
Deliv ery
2
Demonstrat ion
2
Demonstrat ion
2
2
Demonstrat ion
Demonstrat ion
3
Deliv ery
1
Review
2
Review
1
Inspection
Only one person per piggy back pay load will be allowed in launch cent er during launch
1.7
1.8
Shut tle Hit chhiker is capable of 0.3-1.2 m/ s
while the ASAP 5 has t he maximum v elocity
(3 m /s) capability . This and CG dat a will be
required f or Tip-of f and Collision Av oidance
Analy sis
Som e launch vehicles may be able to
prov ide special at titudes and spin at
separation, but commonality def ines this
requirement
This allows t he launch vehicle provider the
maximum leeway in launching wit hout t he
act iv e piggy back pay load in case of
schedule delays
This is only required by Arianespace f or
ASAP 5. Ot her launch v ehicle prov iders
may be more lenient
3
2
Review
2
Review
1
Review
2
Review
3
Review
3
Inspection
2
Review
If required, a Structural Integrit y Verif icat ion Report shall be ready f or rev iew by launch prov ider at
least 13 mont hs bef ore scheduled launch (T - 13 months)
Enginering 176 #5
Sour ce/ Comm ents
Cri t.
Piggy back pay load shall be giv en 0.6 - 1.2 m/s separation v elocit y relative to the launch v ehicle.
Velocit y v ector shall be along pay load longit udinal axis
1.4
1.6
Requi rements
Mission
Piggy back payload insertion orbit shall be dependent on primary pay load orbit
Prim ary pay load orbit shall not be af f ect ed by piggyback orbit
On some m issions, notably STS, there
might be some leeway
See Docum entation sect ion f or documents
that are also due at t he same tim e
This requirement is driv en by STS
Hit chhiker. There may be some leeway f or
piggy back pay load provider
This is required of piggyback pay loads on
Shut tle Hit chhiker. There is some leeway
on Shutt le launch schedules, so delay s
may be negot iated with NASA. This
requirement may be ignored on ot her
launchers
STP-Sat
Requirements
(Some)
System Definition
Requirements & Sys
2.1
Mission Description
Definition go together
2.2
Interface Design
2.2.1
SV-LV Interface
2.2.2
SC-Experiments Interface
2.2.3
Satellite Operations Center (SOC) Interface
3.0
Requirements
3.1
Performance and Mission Requirements
3.2
Design and Construction
3.2.1
Structure and Mechanisms
3.2.2
Mass Properties
3.2.3
Reliability
3.2.4
Environmental Conditions
3.2.4.1 Design Load Factors
3.2.4.2 SV Frequency Requirements
3.2.5
Electromagnetic Compatibility
3.2.6
Contamination Control
3.2.7
Telemetry, Tracking, and Commanding
(TT&C) Subsystem
3.2.7.1
Frequency Allocation
Highly structured
3.2.7.2
Commanding
outline form is
3.2.7.3
Tracking and Ephemeris
3.2.7.4
Telemetry
clearest and
3.2.7.5
Contact Availability
industry standard
3.2.7.6
Link Margin and Data Quality
3.2.7.7
Encryption
Enginering 176 #5
2.0
NB: this is
an excerpt
of the TOC
- the
entire doc
is (or will
be) on the
class FTP
site
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
Enginering 176 #5
∆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)
Enginering 176 #5
∆V = gIspln(R):
Staring at logarithmic reality
3000
∆
V
∆V Performance Samples:
dry mass 50 kg
m
e
t 2000
e
r
s
Isp 300 seconds
p
e 1000
r
Isp 60 seconds
s
e
c
0
0
10
20
30
Propellant mass (kg)
40
50
Staging is an answer...
Enginering 176 #5
Single vs. Two Stage
Assumptions:
• R = M(i)/M(f) = 10
• ∆V required: 10 km/s
• Payload = 100 kg
• Payload =10% Mf
SSTO: 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%
Enginering 176 #5
Enginering 176 #5
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.
Enginering 176 #5
Launch Costs vs. Mission
Costs
• Rationale
• Numbers
– Satellite Cost = Launch Cost
– Scout / Pegasus Payloads
•
•
•
•
•
•
•
ALEXIS + REX:
HETE / SAC-B:
Microsats:
REX / TEX:
Stacksat
8 x Orbcomm
MSTI-2
$24M
$25M
$6M
$6M
$6M
$24M
$14M
– Ariane ASAP class payloads
•
•
•
•
Amsat Oscar
$200k (typ.)
Oscar 13
$200k
4 x Microsats
$200k
Astrid (Kosmos) $1M
– Ariane / Long March
Interstage
• Freja
Enginering 176 #5
$4M
– 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 =
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
Enginering 176 #5
New Options to Orbit
• Candidates
– Aircraft: carry, balloon,
tow
– SSTO: autogyro, Shuttlelike, DC-X, Suborbital
– Sea Launch
• Perspectives
– Jet Aircraft / Ford
(Taurus) costs over
last 40+ years
– Pegasus v. Scout
– “Cheap” Russian rockets
– AF EELV cost goals
(marginal savings)
– Reusable rockets
– Labor cost distortions
– “Cheap” US, Indian,
Spanish, Brazilian,
Chinese or Italian rockets
Enginering 176 #5
– Commercial
Competition: Ariane v.
Long March v. Proton
Space Transportation’s Future
(15 year outlook)
Hint: Nobody lives at the north pole, and launches won’t cost $10/kg
• 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 onboard electric propulsion
• Wildcards: siting and
environmental issues
• Low cost
components ≠
low cost rockets:
hardware vs.
reliability $
Enginering 176 #5
TM
TM
The Next Generation of Microspace
Small Payload
ORbit Transfer
Enginering 176 #5
AeroAstro Proprietary
What is SPORT?
Arianespace
Encounter \ SAIC
TM
Small Payload
ORbit Transfer
Ariane 5
Heavy Launcher
Upper Stage
Propulsion
Microsatellite Going to GTO
(No SPORT)
Enginering 176 #5
Microsatellite Going to Custom Orbit
SPORT
SPORT GTO to LEO Transfer
2
SPORT™
Microsatellite
4
5
3
1
Enginering 176 #5
6
1
2
3
4
5
6
Launch into GTO
Perigee lowering burn
Aerobraking drag near perigee
Apogee reduction with each pass
Perigee raising burn
Final circular orbit
Aerobraking
2
• Highly efficient orbit transfer (over 2 km/s ∆V)
• Rarified atmosphere altitude - minimal heating
• Large deployable increases profile area ( 50)
4
• ~ 200 passes to lower apogee 35,000 km
• Nominal 30 day mission
5
3
1
Enginering 176 #5
6
1
2
3
4
5
6
Launch into GTO
Perigee lowering burn
Aerobraking drag near perigee
Apogee reduction with each pass
Perigee raising burn
Final circular orbit
SPORT Releases Microsatellite
Dispose
SPORT™
Release
Microsatellite in
Custom Orbit
Enginering 176 #5
Aerobraking Performance
Utilizing the aerobraking and propulsion features of SPORT, a
wide range of missions is possible.
Note: Assumes total initial mass of 100 kg.
Enginering 176 #5
™
SPORT performs a variety of
orbit transfer maneuvers
Molniya
to SSO
LEO to MEO
GTO to LEO
L4
GTO
To
GEO
L1
L2
Sun Centered
L5
Enginering 176 #5
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 Capability
Enginering 176 #5
Micro SPORT:
Mini SPORT:
20 kg
60 kg
LEO to MEO Transfer
• Initial Orbit: Polar LEO
3
– 800 km  800 km and 98.6 deg
• Final Orbit: Polar MEO
1
– 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
2
Note: no aerobraking hardware required
Nominal Payload Capability
Micro SPORT:
Mini SPORT:
Enginering 176 #5
50 kg
150 kg
1
2
3
4
Launch into SSO
Perigee burn
Apogee burn
Final circular orbit
4
Direct Transfer Performance
Utilizing 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.
Enginering 176 #5
High Energy Missions
• Initial Orbit: GTO
– 620 km  35,883 km and 7.0 deg
– Launch on Ariane 5 in ASAP Slot
L4
• Final Orbit Options:
– Earth Escape
– Lagrange Point
– Lunar Transfer
– Asteroid Flyby
• SPORT™ Transfer
L1
L5
– V Burn at Perigee
Nominal Payload Capability
Micro SPORT:
Mini SPORT:
Enginering 176 #5
L2
20 kg
60 kg
SPORT Systems
Microsatellite Payload
Payload Interface Ring
Bitsy kernel
• Developed for NASA and USAF
• Includes core satellite capabilities
- Communications
- C&DH
- Power regulation
- G&C
Propulsion System
• Modular per ∆V required
• Simple spin stabilized design
Aerobrake
Batteries
• Provided by proven supplier
• AeroAstro patent pending
• Modular per mission
• Variety of options
based on flight
proven technology
Enginering 176 #5