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Satellite Design Course Spacecraft Configuration Structural Design Preliminary Design Methods
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Transcript of Satellite Design Course Spacecraft Configuration Structural Design Preliminary Design Methods
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Feb 16 2005 ENAE 691 R.Farley NASA/GSFC
Satellite Design CourseSpacecraft Configuration
Structural DesignPreliminary Design Methods
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Feb 16 2005 ENAE 691 R.Farley NASA/GSFC2
Contents Origin of structural requirements pg 3 Spacecraft configuration examples pg 8 Structural configuration examples pg 14 General arrangement drawings pg 25 Launch vehicle interfaces and volumes pg 30 Structural materials pg 33 Subsystem mass estimation techniques pg 38 Vibration primmer pg 51 Developing limit loads for structural design pg 59 Sizing the primary structure pg 68 Structural subsystem mass pg 73
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Feb 16 2005 ENAE 691 R.Farley NASA/GSFC3
Origin of Structural Requirements
This ride not recommended for children
Launch Vehicle
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Feb 16 2005 ENAE 691 R.Farley NASA/GSFC4
Source of Launch Vehicle Loads
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Feb 16 2005 ENAE 691 R.Farley NASA/GSFC5
Typical Loads Time-Line ProfileMax Q
MECOSECO
H2 vehicle to geo-transfer orbit
3rd stage2nd stage
1st stage
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Feb 16 2005 ENAE 691 R.Farley NASA/GSFC6
Design Requirements, Function Launch Vehicle volume, mass, and c.g. Foot print for subsystems Deployment of flexible structures Natural frequency, stowed and deployed Alignment stability (launch shift, thermal, vibration, material shrinkage) Field Of View (FOV) RF and magnetic compatibility Jitter disturbance response Thermal steady state and transient distortions Materials for space environment (out gassing, atomic oxygen)
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Feb 16 2005 ENAE 691 R.Farley NASA/GSFC7
Design Requirements, LoadsTestingTransportationLaunch loads
Steady state accelerations (axial thrust) Steady vibrations (low frequency, resonant burn from solid rocket boosters) Transient vibrations (lift off, transonic, MECO, SECO) Acoustic and random accelerations (lift off, max Q leading to max lateral loads) Shock (payload separation from upper stage adapter) Depressurization (influence on enclosed volumes including blankets inside of instruments)
Orbit Loads spin-up, de-spin, thermal, deployments, maneuvers
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Feb 16 2005 ENAE 691 R.Farley NASA/GSFC8
From the spacecrafts point of view
The sun in the orbit plane changes due to the seasonal change in the suns declination (the tilt in the Earths axis) and orbit plane precession due to the Earths equatorial bulge. At 35o inclination, the precession period is ~ 55 days
Fixed solar array
Articulating solar array, one axis
Articulating solar array, sometimes 2 axis
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Feb 16 2005 ENAE 691 R.Farley NASA/GSFC9
Spacecraft Configuration Examples
EOS aqua
one oar in the water produces an aerodynamic torque that averages to zero per orbit, but reaction wheels must be large enough to absorb in the interim
LEO nadir pointing
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Feb 16 2005 ENAE 691 R.Farley NASA/GSFC10
Configuration Types LEO stellar pointing
Low inclination XTE, HSTHigh inclination WIRE
Articulated solar array Fixed solar array sun-sync orbit
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Feb 16 2005 ENAE 691 R.Farley NASA/GSFC11
Configuration Types GEO nadir pointing
Articulated s/a spin once per 24 hoursTDRSS A
Intelsat V
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Feb 16 2005 ENAE 691 R.Farley NASA/GSFC12
Configuration Types HEO
AXAF-1 CHANDRA required to stare uninterrupted
250m long wire booms, rpm
IMAGE 2000
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Feb 16 2005 ENAE 691 R.Farley NASA/GSFC13
Configuration types misc
MAP at L2Sky surveys in the ecliptic plane
Cassini to Saturn
Lunar ProspectorBody-mounted s/a
Cylindrical, body-mounted solar array
Sun shield
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Feb 16 2005 ENAE 691 R.Farley NASA/GSFC14
Structural Configuration Examples Cylinders and Cones high buckling resistance Box structures hanging electronics and equipment
Triangle structures sometimes needed
Rings transition structures and interfaces
Trusses extending a hard point (picking up point loads) As much as possible, payload connections should be kinematic Skin Frame, Honeycomb Panels, Machined Panels, Extrusions
NOTE: ALWAYS PROVIDE A STIFF AND DIRECT LOAD PATH! AVOID BENDING!
STRUCTURAL JOINTS ARE BEST IN SHEAR!
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Feb 16 2005 ENAE 691 R.Farley NASA/GSFC15
Structural examples
MAGELLAN box, ring, truss
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Feb 16 2005 ENAE 691 R.Farley NASA/GSFC16
Structural examples
OAO cylinder/stringer and decksHESSI ring, truss and deck
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Feb 16 2005 ENAE 691 R.Farley NASA/GSFC17
Structural examples
Archetypical cylinder and box Hybrid one of everything
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Feb 16 2005 ENAE 691 R.Farley NASA/GSFC18
Structural examplesEOS aqua, bus
Hard points created at intersections
Egg crate torque box composite panel bus structure
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Feb 16 2005 ENAE 691 R.Farley NASA/GSFC19
Structural examples
COBE
Delta II version
COBE
STS version
5000 kg !
2171 kg !
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Feb 16 2005 ENAE 691 R.Farley NASA/GSFC20
Structural examplesTRIANA L1 orbit
Boxes mounted to outer panels to radiate heat
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Feb 16 2005 ENAE 691 R.Farley NASA/GSFC21
Anatomy of s/c
Instrument module
Bus module
Propulsion module
Launch Vehicle adapter
Interface isolation structure
Axial viewing, telescope type
Astronomy Mission
Solid kick motor, equatorial mount
Station-keeping hydrazine, polar mountReaction wheels
Deck-mounted or wall-mounted boxes
Kinematic Flexure mounts to remove enforced displacement loads
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Feb 16 2005 ENAE 691 R.Farley NASA/GSFC22
Modular AssemblyInstrument Module optics, detectorBus Module (house keeping)Propulsion Module
Modules allow separate organizations, procurements, building and testing schedules. It all comes together at observatory integration and test (I&T)
Interface control between modules is very important: structural, electrical, thermal
AXAF
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Feb 16 2005 ENAE 691 R.Farley NASA/GSFC23
Attaching distortion-sensitive components
2,2,2
Bi-pod legs, tangential flexures
Breathes from center point
Rod flexures arranged in 3,2,1
Breathes from point 3
1,1,1,1,1,1
32
1
1
2
3
Note: vee points towards cone3,2,1
Ball-in-cone, Ball-in-vee, Ball on flat
Breathes from point 3
JPL perfect joint for GALEX primary mirror
Hexapod arrangement
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Feb 16 2005 ENAE 691 R.Farley NASA/GSFC24
General Arrangement Drawings
Stowed configuration in Launch Vehicle Transition orbit configuration Final on-station deployed configuration Include Field of View (FOV) for instruments and thermal
radiators, and communication antennas, and attitude control sensors
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Feb 16 2005 ENAE 691 R.Farley NASA/GSFC25
3-view layout, on-orbit configuration
Overall dimensions and field of views (FOV)
Antennas showing field of regard (FOR)
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Feb 16 2005 ENAE 691 R.Farley NASA/GSFC26
3-view layout, launch configuration
Make note of protrusions into payload envelope
Omni antennasHigh gain antenna, stowed
Star trackers Payload Adapter Fitting (PAF)
Solar array panels array drive
Torquer bars
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Feb 16 2005 ENAE 691 R.Farley NASA/GSFC27
Launch Vehicle Interfaces and VolumesPayload Fairings
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Feb 16 2005 ENAE 691 R.Farley NASA/GSFC28
Pegasus Interface
Coupled loads analysis necessary
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Feb 16 2005 ENAE 691 R.Farley NASA/GSFC29
Payload Envelope
If frequency requirements are not met, or for protrusions outside the designated envelope, Coupled Loads Analysis (CLA) are required to qualify the design, in cooperation with the LV engineers.
For many vehicles, if the spacecraft meets minimum lateral frequency requirements, then the static envelope accounts for fairing dynamic motions.
STATIC ENVELOPE; the space that the payload must fit inside of when integrated to the vehicle
DYNAMIC ENVELOPE; the space that the payload must stay in during launch to account for all payload deflections
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Feb 16 2005 ENAE 691 R.Farley NASA/GSFC30
Launch Vehicle Payload Adapters
6019 3-point adapter 6915 4-point adapter
V-band clamp, or Marmon clamp-band
Clamp-band
Shear Lip
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Feb 16 2005 ENAE 691 R.Farley NASA/GSFC31
Structural MaterialsMaterial
(kg/m3)E (GPa)
Fty(MPa)
E/ Fty/ (m/m K)
(W/m K)
Aluminum6061-T67075-T651
28002700
6871
276503
2426
98.6186.3
23.623.4
167130
MagnesiumAZ31B 1700 45 220 26 129.4 26 79
Titanium6Al-4V 4400 110 825 25 187.5 9 7.5
BerylliumS 65 AS R 200E
2000-
304-
207345
151-
103.5-
11.5-
170
FerrousINVAR 36AM 350304L annealed4130 steel
8082770078007833
150200193200
62010341701123
18.5262525
76.7134.321.8143
1.6611.917.212.5
1440-601648
Heat resistantNon-magneticA286Inconel
600Inconel
718
794484148220
200206203
5852061034
252425
73.624.5125.7
16.4-23.0
12-12
Metallic Guide
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Feb 16 2005 ENAE 691 R.Farley NASA/GSFC32
Materials Guide Definitions
mass density, kg/m3
E Youngs modulus, Pascals
Fty material allowable yield strength Note: Pa = Pascal = N/m2
E /
specific stiffness, the ratio of stiffness to density
Fty /
specific strength, the ratio of strength to density
coefficient of thermal expansion CTE
coefficient of thermal conductivity
Material Usage Conclusions: USE ALUMINUM WHEN YOU CAN!!!Aluminum 7075 and Titanium 6Al-4V have the greatest strength to mass ratio
Beryllium has the greatest stiffness to mass ratio and high damping
4130 Steel has the greatest yield strength
INVAR has the lowest coefficient of thermal expansion, but difficult to process
Titanium has the lowest thermal conductivity, good for metallic isolators
(expensive, toxic to machine, brittle)
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Feb 16 2005 ENAE 691 R.Farley NASA/GSFC33
Metallic Materials Usage Guide
High strength to weight, good machining, low cost and available
High strength to weight, low CTE, low thermal conductivity, good at high temperatures
High stiffness, strength, low cost, weldable
High stiffness, strength at high temperatures, oxidation resistance and non-magnetic
Very high stiffness to weight, low CTE
Aluminum
Titanium
Steel
Heat-resistant Steel
Beryllium
Poor galling resistance, high CTE
Expensive, difficult to machine
Heavy, magnetic, oxidizes if not stainless steel. Stainless galls easily
Heavy, difficult to machine
Expensive, brittle, toxic to machine
Truss structure, skins, stringers, brackets, face sheets
Attach fittings for composites, thermal isolators, flexures
Fasteners, threaded parts, bearings and gears
Fasteners, high temperature parts
Mirrors, stiffness critical parts
Advantages Disadvantages Applications
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Feb 16 2005 ENAE 691 R.Farley NASA/GSFC34
Subsystem Mass Estimation TechniquesPreliminary Design Estimates for Instrument Mass
Approximate instrument mass densities, kg / m3
Spectrometers ~ 250
Mass spectrometers ~ 800
Synthetic aperture radar ~ 32
Rain radars ~ 150 thickness / diameter ~ 0.2
Cameras ~ 500 Small telescopes w/ camera ~ 325
Small instruments ~ 1000
Scaling Laws: If a smaller instrument exists as a model, then if SF is the linear dimension scale factor
Area proportional to SF2 Mass proportional to SF3
Area inertia proportional to SF4 Mass inertia proportional to SF5
BEWARE THE SQUARE-CUBE LAW! STRESS WILL INCREASE WITH SF!
Frequency proportional to 1/ SF Stress proportional to SF
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Feb 16 2005 ENAE 691 R.Farley NASA/GSFC35
Typical List of Boxes, Bus ComponentsACS
Reaction wheels Torquer bars Nutation damper Star trackers Inertial reference unit Earth scanner Digital sun sensor Coarse sun sensor Magnetometer ACE electrical box
Communication
S-band omni antenna S-band transponder X-band omni antenna X-band transmitter Parabolic dish reflector 2-axis gimbal Gimbal electronics Diplexers, RF switches Band reject filters Coaxial cable
Power
Batteries Solar array panels Articulation mechanisms Articulation electronics Array diode box Shunt dissipaters Power Supply Elec. Battery a/c ducting
Mechanical
Primary structure Deployment mechanisms Fittings, brackets, struts, equipment decks, cowling, hardware Payload Adapter Fitting
Propulsion
Propulsion tanks Pressurant tanks Thrusters Pressure sensors Filters Fill / drain valve Isolator valves Tubing
Thermal
Radiators Louvers Heat pipes Blankets Heaters Heat straps Sun shield Cryogenic pumps Cryostats
Electrical
C&DH box Wire harness
Instrument electronics
Instrument harness
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Feb 16 2005 ENAE 691 R.Farley NASA/GSFC36
Mass Estimation, mass fractions Mi /MdrySome Typical Mass Fractions for Preliminary Design
Payload Structure Power Electricalharness
ACS Thermal C&DH Comm Propulsion, dry
LEOnadir(GPM)
37 % 24 % 13 %Fixedarrays
7 % 6 % 4 % 1 % 3 % 5 %26 % w/fuel
LEOstellar(COBE)
52 % 14 % 12 %spins at 1 rpm
8 % 8 % 2 % 3 % 1 % 0 %
GEOnadir(DSP 15)
37 % 22 % 20 %spins at6 rpm
7 % 6 % 0.5 % 2 % 2 % 2 %
Mass fractions as percentage of total spacecraft observatory mass, less fuel
COBE with 52% payload fraction is unusually high and not representativeACS is Attitude Control System, C&DH is Command and Data Handling, Comm is Communication subsystem
LEO (Pegasus class)nadir(FireSat)
13-20 % 30 % 17 %Articulat
ing arrays
7 % 14 % 1 % 3 % 15 %2-axis gimball
0 %
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Feb 16 2005 ENAE 691 R.Farley NASA/GSFC37
Mass Estimation, RefinementsThe largest contributors to mass are (neglecting the instrument payload)
Structural subsystem (primary, secondary)
Propulsion fuel mass, if required
Power subsystem
The following slides will show the cheat-sheet for making preliminary estimates on some of these subsystems
Remember to target a mass margin of ~20% when compared to the throw weight of the launch vehicle and payload adapter capability. There must be room for growth, because evolving from the cartoon to the hardware, it always grows!
Define ratio mi = Mi / MdryMdry = mass of spacecraft less fuel
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Feb 16 2005 ENAE 691 R.Farley NASA/GSFC38
Mass Estimation with Mass Ratios
mpayload ~ 0.4(large sat) ~0.2(small sat) for a first guess ratio
Mdry = Mpayload / mpayload
Mdry = Mpayload + Mpower + MPropDry + Mstruct + Melec + MCDH + MACS + Mcomm + Mtherm
Dry mass = sum of subsystem masses without fuel
Mdry = Mpayload + Mdry (mpower + mPropDry + mstruct + melec + mCDH + mACS + mcomm + mtherm )
1st GUESS
2nd GUESS
Mdry = Mpayload + Mpower + MPropDry + Mdry (mstruct + melec + mCDH + mACS + mcomm + mtherm )
3rd GUESS, more refined
Mwet = Mdry + Mfuel Total wet mass, observatory massMlaunch = Mwet + payload adapter fitting Launch mass
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Feb 16 2005 ENAE 691 R.Farley NASA/GSFC39
Mass Ratios, Generalized Formula
M dry = known
M i
1
unknown
m j
known
M i = M payload M power M PropDry
Generalized FormulaSum of known masses / (1-sum of unknown masses as ratios)
Typical knowns + calculated
Typical unknownsm structure m elec m CDH
unknown
m j =+ m ACS m comm m therm
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Feb 16 2005 ENAE 691 R.Farley NASA/GSFC40
Mass Estimation exampleLow Earth Orbit Earth-Observing Mission
Known or calculated masses:
Mpayload = 750 kg
Mpower = 200 kg
MPropDry = 40 kg
Will use these ratios:
melec = 0.07
mCDH = 0.01
mtherm = 0.04
mcomm = 0.03
mACS = 0.06
mstruct = 0.20
known
M i = 750 200 40 = 990
0.20 0.07 0.01 0.41
unknown
m j = =+ 0.06 0.03 0.04
M dry =990
1 0.41( )= 1678 kg
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Feb 16 2005 ENAE 691 R.Farley NASA/GSFC41
Mass Estimation, Power Subsystem
Givens:
ALT circular orbit altitude, km
Poa required orbital average power, watts
AF area factor, if articulated arrays AF = 1, if omni- directional in one plane, AF = 3.14, spherical coverage AF = 4
cell standard cell efficiency, 0.145 silicon, 0.18 gallium, 0.25 multi-junction
If the required orbital average power is not settled, then estimate with: Poa ~ PREQpayload + 0.5 (MDry Mpayload ) watts, mass in kg
Dry bus mass, kg, ~ 0.5 watts/kg
Power for payload instruments and associated electronics ~ 1 watt/kg * Mpayload
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Feb 16 2005 ENAE 691 R.Farley NASA/GSFC42
Mass Estimation, Power Subsystem
Total solar array panel area considering energy balance going thru eclipse and geometric area factors, m2
Maximum eclipse-to- daylight time ratio, estimate good for 300
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Feb 16 2005 ENAE 691 R.Farley NASA/GSFC43
Mass Estimation, Power SubsystemMass of cells, insulation, wires, terminal boards:
Melec ~ 3.8 x
ASA for multi-junction cells kg
Melec ~ 3.4 x
ASA for silicon cells kg
Mass of honeycomb substrates:
Msubstrates ~ 2.5 x
ASA for aluminum panels kg
Msubstrates ~ 2.0 x
ASA for composite panels kg
Mass of solar array panels, electrical and structure:
MPANELS = Melec + Msubstrates kg
Since multi junction cells usually go on composite substrates, and silicon cells went on aluminum substrates, it all comes out in the wash:
MPANELS ~ 5.85 x
ASA kg
93% packing efficiency
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Feb 16 2005 ENAE 691 R.Farley NASA/GSFC44
Mass Estimation, Power Subsystem
Mbattery = 0.06 Poa Battery mass, kg Ni-Cads
MPSE = 0.04 Poa Power systems electronics, kg
MSAD = 0.33 Mpanels Solar array drives and electronics, kg
MSAdeploy = 0.27 Mpanels Solar array retention and deployment mechanisms, kg
Mpower = Mpanels + MSAD + MSAdeploy + Mbattery + MPSE
Total power subsystem mass estimate, kg
(this can vary greatly due to the peak power input and charging profile allowed. High noon power input can be enormous sometimes, depending on orbit inclination and drag reduction techniques)
Tracking array
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Feb 16 2005 ENAE 691 R.Farley NASA/GSFC45
Power System Mass Example, givensALT = 700 km circular orbit altitude, sun synchronous
AF = 1 area factor
cell = 0.25 standard cell efficiencyMdry = 1000 kg
Mpayload = 350 kg (35% of dry mass)
Ppayload = 350 watts (1 watt/kg * 350kg) estimate
Pbus = 325 watts (0.5 watts/kg * (1000-350)) estimate
Poa = Ppayload + Pbus = 350 + 325 = 675 watts estimated orbital average power
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Feb 16 2005 ENAE 691 R.Farley NASA/GSFC46
Power System Mass Example, calculationsEtoD = 0.576 6.069 10 4. ALT. 1.768 10 7. ALT2. 0.238=
Area SA = AF P oa.1
0.9EtoD0.54
997 cell.. 4.203=
M panels = 5.85 Area SA. 24.6=
M battery = 0.06 P oa. 40.5=
M PSE = 0.04 P oa. 27=
M SAD = 0.33 M panels. 8.1=
M SAdeploy = 0.27 M panels. 6.6=
M power = M panels M SAD M SAdeploy M battery M PSE 106.8=
Sun synchronous
kg
Total panel area m2
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Feb 16 2005 ENAE 691 R.Farley NASA/GSFC47
Mass Estimation, Propulsion Subsystem
Givens:
Isp fuel specific thrust, seconds (227 for hydrazine, 307 bi-prop)
V deltaV required for maneuvers during the mission, m/sMdry total spacecraft observatory mass, dry of fuel, kg
Isp is the amount of kick that the fuel has to offer per kg
dMfuel /dt = Thrust / (g*Isp) kg per second
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Feb 16 2005 ENAE 691 R.Farley NASA/GSFC48
Mass Estimation, Propulsion Subsystem
M fuel = M dry e
VIsp g. 1.
If the fuel is to be saved for a de-orbit maneuver:
Or expressed as a ratio: =M fuelM dry e
VIsp g. 1
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Feb 16 2005 ENAE 691 R.Farley NASA/GSFC49
Mass Estimation, Propulsion SubsystemApproximate V to de-orbit to a 50km perigee:From 400km circular ~ 100 m/s
From 500km circular ~ 130 m/s
From 600km circular ~ 160 m/s
From 700km circular ~ 185 m/s
M fuel M fuelDRAG M fuelMANV Total fuel mass, kg
M PropDry 0.25 M fuel. Propulsion system mass, dry, kgtank, thrusters, lines, etc...
M wet M dry M fuel Total spacecraft observatory mass with fuel, kg
Remember, the dry mass estimate now includes the mass of the dry prop system components.
This means iterating with a better dry mass estimate for a better fuel calculation.
If the thrusters are too small for one large V de-orbit maneuver, then 2-3 times the calculated fuel maybe required (TRMM experience)
Add 10% for ullage
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Feb 16 2005 ENAE 691 R.Farley NASA/GSFC50
Fuel Estimate Example
M dry e
VIsp g. 1. = 1000 e
185227 9.807. 1. 86.7=
Orbit is a 700km circular, so V to de-orbit ~185 m/sFuel is hydrazine, so Isp = 227 seconds
g = 9.807 m/s2
Mdry = 1000 kg
Mfuel =
M PropDry = 0.25 M fuel. 21.7= kg
kg
MS/C wet = Mfuel + Mdry = 1087 kg
Fuel tank, thrusters, fuel lines, valves, etc
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Feb 16 2005 ENAE 691 R.Farley NASA/GSFC51
Vibration Primmer A vibrating structure can be thought of as the
superposition of many mode shapes Each mode has a particular natural resonant
frequency and distortion shape If the resonant frequencies are sufficiently
separated, then the structural response can be estimated by treating each mode as a single-degree-of-freedom (sdof) system
3 usual flavors of vibration:
Harmonic Random Impulsive
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Feb 16 2005 ENAE 691 R.Farley NASA/GSFC52
Harmonic (sinusoidal) Vibration, sdof
Hz
0.05 typical
Natural frequency of vibration for a single degree of freedom system
Recall:
2
f =
Circular frequency, radians / s
Transfer function
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Feb 16 2005 ENAE 691 R.Farley NASA/GSFC53
Harmonic Vibration, sdof cont
Typical structural values for Q: 10 to 20
For small amplitude (jitter), or cryo temperatures: Q >100
Resonant Load Acceleration = input acceleration x
Q
f is the applied input sinusoidal frequency, Hz
fn is the natural resonant frequency
Harmonic output acceleration = Harmonic input acceleration x
T
Gain function
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Feb 16 2005 ENAE 691 R.Farley NASA/GSFC54
Natural Frequency Estimation evenly distributed mass and stiffness
E is the Youngs Modulus of Elasticity of the structural material
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Feb 16 2005 ENAE 691 R.Farley NASA/GSFC55
Natural Frequency Estimation discrete and distributed mass and stiffness
E is the Youngs Modulus of Elasticity of the structural material
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Feb 16 2005 ENAE 691 R.Farley NASA/GSFC56
Base-Driven Random Vibration Random loads affect
mostly smaller, high frequency components, not the overall spacecraft structure
Input spectrum So described in power spectral density g2 per Hz
Miles equation gives a statistical equivalent peak g load for a single degree of freedom system, 68% of the time (1). For 99.73% confidence (3), multiply by 3.
fn is the natural frequency of the system, Q is the resonant amplification, and So is the input acceleration power spectral density, g2/Hz
So
Mechanically conducted random and Acoustically conducted random
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Feb 16 2005 ENAE 691 R.Farley NASA/GSFC57
Developing Limit Loads for Structural Design
Quasi-Static loads Linear and rotational accelerations + dynamicDynamic Loads included in Quasi-Static: Harmonic vibration (sinusoidal) Random vibration (mostly of acoustic origin) Vibro-acoustic for light panels
Thermal loadsDont forget, there are payload mass vs. c.g. height limitations for the launch vehicles payload adapter fitting Limit loads do not have
stress factors of safety incorporated yet
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Feb 16 2005 ENAE 691 R.Farley NASA/GSFC58
Launch Loads and Fundamental FrequencyLaunch Vehicle Axial Load Factor Naxial
, gs Lateral Load Factor Nlateral
, gs
Delta IVsteady statedynamic
6.5+/-
1.0
27Hz min+/-
2.0+/-
0.7
10Hz min
Atlas IIsteady statedynamic
5.5+/-
2.0
15Hz min+/-
0.4+/-
1.2
10Hz min
H-IIsteady statedynamic
5.2+/-
5.0 30Hz min2.8+/-
2.0 10Hz min
Ariane
Vsteady statedynamic
4.6+/-
less than 1.0
18Hz min+/-
0.25+/-
0.8
10Hz min
Delta IIsteady statedynamic
6.7 for 2000kg+/-
1.0 35Hz min+/-
3.0+/-
0.7 15Hz min
Note: the static and dynamic loads do not occur at the same time (usually)
Beware 32Hz MECO-POGO
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Feb 16 2005 ENAE 691 R.Farley NASA/GSFC59
Pegasus LoadsFor Pegasus, quasi-static loads are equivalent to low frequency dynamic loads
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Feb 16 2005 ENAE 691 R.Farley NASA/GSFC60
Pegasus Loads cont
First lateral bending frequency > 20 Hz
To simplify, boil down to 3 load case combinations to analyze:
Drop transient z ~ 6.5g lateral
Stage burnout 10.5g axial, 1.7g lateral
Aerodynamic pull-up 4.7g axial 3.5g lateral
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Feb 16 2005 ENAE 691 R.Farley NASA/GSFC61
Loads for primary structure, exampleIf the structure meets minimum frequency requirements from the launch vehicle, then the low frequency sinusoidal environment is enveloped in the limit loads. Secondary structures with low natural frequencies may couple in, however, and
The structural analyst will determine which load case produces the greatest combined axial-bending stress in the structure (I, A, mass and c.g. height)
Reaction loads
should be analyzed separately.
Lcg
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Feb 16 2005 ENAE 691 R.Farley NASA/GSFC62
Loads example, cont2 major load case combinations:
a) 0.5g lateral + 6.5g axial
b) 2.0g lateral + 3.5g axial
Case a) load combination:
Axial load = Mwet *Naxial *g = 2000*6.5*9.807 = 127491 N
Lateral load = Mwet *Nlateral *g = 2000*0.5*9.807 = 9807 N
Moment = Lateral load * Lcg = 9807*1.5 = 14710.5 N-m
Given: Mwet = 2000 kg
g = 9.807m/s2
Lcg = 1.5m
Case b) load combination:
Axial load = Mwet *Naxial *g = 2000*3.5*9.807 = 68649 N
Lateral load = Mwet *Nlateral *g = 2000*2.0*9.807 = 39228 N
Moment = Lateral load * Lcg = 39228*1.5 = 58842 N-m
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Feb 16 2005 ENAE 691 R.Farley NASA/GSFC63
Component Loads, Mass-Acceleration CurveJPL acceleration curve for component sizing
gloadi
massi1 10 100 1 103
1
10
100Mass Acceleration Curve
Mass kg
A
c
c
e
l
e
r
a
t
i
o
n
g
'
s
This curve envelopes limit loads for small components under 500 kg
Apply acceleration load separately in critical direction
Add static 2.5 g in launch vehicle thrust direction
Applicable Launch Vehicles:
STS
Titan
Atlas
Delta
Ariane
H2
Proton
Scout
Simplified design curve for components on appendage-like structures under 80 Hz fundamental frequency
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Feb 16 2005 ENAE 691 R.Farley NASA/GSFC64
Sizing the Primary Structure rigidity, strength and stability
Factors of safety NASA / INDUSTRY, metallic structures
Factors of safety Verification by Test Verification by Analysis
FS yield 1.25 / 1.10 2.0 / 1.6
FS ultimate 1.4 / 1.25 (1.5 Pegasus) 2.6 / 2.0 (2.25 Pegasus)
Factors of safety for buckling (stability) elements ~ FS buckling = 1.4
(stability very dependent on boundary conditions.so watch out!)
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Feb 16 2005 ENAE 691 R.Farley NASA/GSFC65
Primary Structure Simplified Model
For example, in many cases, the primary structure is some form of cylinder
E is the Youngs Modulus of Elasticity of the structural material
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Feb 16 2005 ENAE 691 R.Farley NASA/GSFC66
Sizing for Rigidity (frequency)
Working the equations backwards.
When given frequency requirements from the launch vehicle users guide for 1st major axial frequency and 1st major lateral frequency, faxial , flateral :
Material modulus E times cross sectional area A
Material modulus E times bending inertia I
For a thin-walled cylinder:
I
=
R3 t , or t = I
/ (
R3)
A = 2
R t or t = A / (2
R) Determine the driving requirement resulting in the thickest wall t. Recalculate A and I with the chosen t.
Select material for E, usually aluminum
e.g.: 7075-T6 E = 71 x 109 N/m2
Calculate for a 10% 15% frequency marginTapering thickness will drop frequency 5% to 12%, but greatly reduce structural mass
AXIAL STIFFNESS
BENDING STIFFNESS
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Feb 16 2005 ENAE 691 R.Farley NASA/GSFC67
Sizing for Strength
Plateraldes = (M + Mtip ) g NlimitL Lateral Design Load, N
Paxialdes = (M + Mtip ) g NlimitA Axial Design Load, N
Momentdes = Plateraldes * Lcg Moment Design Load, N-m
Design Loads using limit loads and factors of safety
Recalling from mechanics of materials:
axial stress = P/A, bending stress = Moment*R/I
(in a cylinder, the max shear and max compressive stress occur in different areas and so for preliminary design shear is not considered)
Max stress max = Paxialdes / A + (Plateraldes Lcg R) / IMargin of safety MS = {allowable / (FS x
max )} 1 0 < MS acceptableFor 7075-T6 aluminum, the yield allowable , allowable = 503 x 106 N/m2
With less stress the higher up, the more tapered the structure can be, saving mass
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Feb 16 2005 ENAE 691 R.Farley NASA/GSFC68
Sizing for Structural StabilityDetermine the critical buckling stress for the cylinder
In the general case of a cone:
Allowable Critical Buckling Stress
Compare critical stress to the maximum stress as calculated in the previous slide. Update the max stress if a new thickness is required.
Margin of safety MS = {CR /(FSbuckling x
max )} 1 0 < MS acceptableCheck top and bottom of cone: max , I, moment arm will be different Sheet and stringer construction will save ~
25% mass
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Feb 16 2005 ENAE 691 R.Farley NASA/GSFC69
Structural Subsystem MassFor all 3 cases of stiffness, strength, and stability, optimization calls for tapering of the structure.
The frequency may drop between 5% to 12% with tapering
But the primary structural mass savings may be 25% to 35% - a good trade
Secondary structure (brackets, truss points, interfaces.) may equal or exceed the primary structure. An efficient structure, assume secondary structure = 1.0 x primary structure. A typical structure, assume 1.5.
So, if the mass calculated for the un-optimized constant-wall thickness cylinder (primary structure) is MCYL , then the typical structure (primary + secondary):
MSTRUCTURE ~ (2.0 to 3.5) x
MCYL
This number can vary significantly
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Feb 16 2005 ENAE 691 R.Farley NASA/GSFC70
Backup Charts, Extras
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Feb 16 2005 ENAE 691 R.Farley NASA/GSFC71
Anatomy of s/cTransverse viewing, Earth observing type
Launch Vehicle I/F
Drag make-up Propulsion module
FOV
Instruments
Cylinder-in-boxEgg-crate extension
Heat of boxes radiating outward
Hydrazine tank equatorial mount on a skirt with a parallel load path ! The primary structure barely knows its there.
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Feb 16 2005 ENAE 691 R.Farley NASA/GSFC72
Spacecraft Drawing in Launch Vehicle
XTE in Delta II 10 fairingFairing access port for the batteries
Pre-launch electrical access red-tag item
Launch Vehicle electrical interface
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Feb 16 2005 ENAE 691 R.Farley NASA/GSFC73
Drawing of deployment phase
EOS aqua
Pantograph deployment mechanism
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Feb 16 2005 ENAE 691 R.Farley NASA/GSFC74
Composite Materials
Pros
Lightweight (Strength to Weight ratio)
Ability to tailor CTE
High Strength
Good conductivity in plane
Thermal property variation possible (K1100)
Low distortion due to zero CTE possible
Ability to coat with substances (SiO)
ConsCostly
Tooling more exotic, expensive/specialized tooling (higher rpms, diamond tipped)
Electrical bonding a problemSome types of joints are more difficult to
produce/design
Fiber print through (whiskers)
Upper temperature limit (Gel temperature)
Moisture absorption / desorption / distortion
provided by Jeff Stewart
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Feb 16 2005 ENAE 691 R.Farley NASA/GSFC75
Composite Material PropertiesGraphite Fiber Reinforced Plastic (GFRP) density ~ 1800 kg/m3
If aluminum foil layers are added to create a quasi-isotropic zero coefficient of thermal expansion (CTE < 0.1 x10-6 per Co) then density ~ 2225 kg/m3
Aluminum foil layers are used to reduce mechanical shrinkage due to de- sorption / outgassing of water from the fibers and matrix (adhesive,ie. epoxy)
NOTE! A single pin hole in the aluminum foil will allow water de-sorption and shrinkage. This strategy is not one to trust
Cynate esters are less hydroscopic than epoxies
Shrinkage of a graphite-epoxy optical metering structure due to de-sorption may be described as an asymptotic exponential (HST data):
Shrinkage ~ 27 (1 e - 0.00113D) microns per meter of length, where D is the number of days in orbit.
Shrinkage D( ) 27 1 e 0.00113D..
Shrinkage D( )
D0 1000 2000 3000 4000 5000
0
20
40Graphite-Epoxy Shrinkage in Vacuum
Days
M
i
c
r
o
n
s
p
e
r
M
e
t
e
r
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Feb 16 2005 ENAE 691 R.Farley NASA/GSFC76
Mass Estimation, Propulsion Subsystem
Givens:
Isp fuel specific thrust, seconds (227 for hydrazine, 307 bi-prop)
V deltaV required for maneuvers during the mission, m/sTM mission time in orbit, years
YSM years since last solar maximum, 0 to 11 years
AP projected area in the velocity direction, orbital average, m2
ALT circular orbit flight altitude, km
Mdry total spacecraft observatory mass, dry of fuel, kg
Area SAoa Area SA
1 cos
EtoD 1
.Solar array orbital average projected drag area for a tracking solar array that feathers during eclipse
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Feb 16 2005 ENAE 691 R.Farley NASA/GSFC77
Mass Estimation, Propulsion Subsystem
If the mission requires altitude control, this is the approximate fuel mass for drag over the mission life
Note: g = 9.807 m/s2
Densitymax = 4.18 x
10-9 x
e(-0.0136 ALT)
DF = 1 0.9 sin [
x
YSM / 11]
Densityatm = DF x
Densitymax
Approximate maximum atmospheric density at the altitude ALT (km), kg / m3
good for 300 < ALT < 700 km
Density factor, influenced by the 11 year solar maximum cycle (sin() argument in radians)
Corrected atmospheric density, kg / m3
V circular3.986 1011.6371 ALT( )
C D 2.2 Assumed Drag coefficient
Drag12Density atm. V circular2. C D. A P.
M fuelDRAGDrag TM. 31.536. 106.
Isp g.
Circular orbit velocity, m/s
Drag force, Newtons
Fuel mass for drag make-up, kg
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Feb 16 2005 ENAE 691 R.Farley NASA/GSFC78
Mass Estimation, Propulsion Subsystem
Maneuvering fuel mass, kgbeginning of mission caseM fuelMANV M dry M fuelDRAG e
VIsp g.( ) 1.
Maneuvering fuel mass, kgend of mission case (de-orbit)M fuelMANV M dry e
VIsp g.( ) 1.
Fuel mass for maneuveringIf maneuvers are conducted at the beginning of the mission (attaining proper orbit):
If the fuel is to be saved for a de-orbit maneuver:
Or expressed as a ratio:M fuelMANV
M dry = e
VIsp g.( ) 1
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Feb 16 2005 ENAE 691 R.Farley NASA/GSFC79
Deployable Boom, equations for impact torque
Maximum Impact torque at lock-in, N-m
Hinge point o
Io Rotational mass inertia about point o, kg-m2
Dynamic system is critically damped
d/dt can just be the velocity at time of impact if not critically damped
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Feb 16 2005 ENAE 691 R.Farley NASA/GSFC80
Static Beam DeflectionsFor Quick Hand Calculations, these are the most common and useful
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Feb 16 2005 ENAE 691 R.Farley NASA/GSFC81
Rigid-body accelerations
Linear force F = m x a = m x g x Nfactor N is the load factor in gs Low frequency sinusoidal below first natural
frequency will produce near-static accelerationa = A x (2
f)2 where f is the driving frequency and A
is the amplitude of sinusoidal motion Rotational torque Q = I x alpha Centrifugal force Fc = m x r x 2
x = A sin(t)v = A
cos(t)a = -A2 sin(t)
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Feb 16 2005 ENAE 691 R.Farley NASA/GSFC82
Combining the loads to form quasi-static levels
Limit Load = Static + dynamic + resonant + random
Note: The maximum values for each usually occur at different times in the launch environment, luckily.
The primary structure will have a different limit load than attached components.
Solar arrays and other low area-density exposed components will react to vibro-acoustic loads.
(low frequency)
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Feb 16 2005 ENAE 691 R.Farley NASA/GSFC83
Loads for a component, exampleMpayload = 500 kg
Mkickmotor = 50 kg dry mass
Tstatic = 30000 N
Tdynamic = +/- 10% Tstatic at 150 Hz
m = 1 kg
fnaxial = 145 Hz with Q = 15
fnlateral = 50 Hz with Q = 15
L = 0.5m
R = 1m
= 10.5 rad/s (100 rpm)Random input So = 0.015 g2 per Hz
Antenna boom component
The axial frequency is sufficiently close to the driven dynamic frequency that we can consider the axial mode to be in resonance.
(resonant burn, chuffing)
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Feb 16 2005 ENAE 691 R.Farley NASA/GSFC84
Spinning upper stage example, cont
y
Axial-to-lateral coupling
= 2
fnLateral
L R
Length L
Radius R
Deflection y
Lateral circular bending frequency, rad/s
y = R (
/ )2[ 1 (
/ )2 ]
Deflection y:Paxial
Plateral
Axial-to-lateral coupling AtoL = y / L
Equivalent additional lateral load = AtoL x
Paxial
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Feb 16 2005 ENAE 691 R.Farley NASA/GSFC85
Loads for a component, cont
Static axial acceleration GstaticA = Tstatic / g (Mpayload + Mkickmotor ) gsDynamic axial acceleration GdynA = Tdynamic x
Q / g (Mpayload + Mkickmotor ) gsRandom axial acceleration GrmdA = 3 sqr[0.5
fnaxial Q So] gs
Axial Limit Load Factor, gs, Naxial = GstaticA + GdynA + GrmdA gsAxial Limit Load, Newtons, Paxial = g x
m x
NaxialL
Static lateral acceleration GstaticL = (R+y) 2 / g gsRandom lateral acceleration GrmdL = AtoL x
3 sqr[0.5
fnlateral Q So] gsDynamic lateral acceleration GdynL ~ 0
Lateral Limit Load Factor, gs, Nlateral = GstaticL + GdynL + GrmdL gsLateral Limit Load, Newtons Plateral = g x
m x
NlateralMoment at boom base Moment = L x Plateral + y x Paxial
Satellite Design Course ContentsOrigin of Structural RequirementsSource of Launch Vehicle LoadsTypical Loads Time-Line ProfileDesign Requirements, FunctionDesign Requirements, LoadsFrom the spacecrafts point of viewSpacecraft Configuration ExamplesConfiguration Types LEO stellar pointingConfiguration Types GEO nadir pointingConfiguration Types HEOConfiguration types miscStructural Configuration ExamplesStructural examplesStructural examplesStructural examplesStructural examples Structural examplesStructural examplesAnatomy of s/cModular AssemblyAttaching distortion-sensitive componentsGeneral Arrangement Drawings3-view layout, on-orbit configuration3-view layout, launch configurationLaunch Vehicle Interfaces and VolumesPegasus InterfacePayload EnvelopeLaunch Vehicle Payload AdaptersStructural MaterialsMaterials Guide DefinitionsMetallic Materials Usage GuideSubsystem Mass Estimation TechniquesTypical List of Boxes, Bus ComponentsMass Estimation, mass fractions Mi/MdryMass Estimation, RefinementsMass Estimation with Mass RatiosMass Ratios, Generalized FormulaMass Estimation exampleMass Estimation, Power SubsystemMass Estimation, Power SubsystemMass Estimation, Power SubsystemMass Estimation, Power SubsystemPower System Mass Example, givensPower System Mass Example, calculationsMass Estimation, Propulsion SubsystemMass Estimation, Propulsion SubsystemMass Estimation, Propulsion SubsystemFuel Estimate ExampleVibration PrimmerHarmonic (sinusoidal) Vibration, sdofHarmonic Vibration, sdof contNatural Frequency Estimationevenly distributed mass and stiffnessNatural Frequency Estimationdiscrete and distributed mass and stiffnessBase-Driven Random VibrationDeveloping Limit Loads for Structural DesignLaunch Loads and Fundamental FrequencyPegasus LoadsPegasus Loads contLoads for primary structure, exampleLoads example, contComponent Loads, Mass-Acceleration CurveSizing the Primary Structure rigidity, strength and stabilityPrimary Structure Simplified ModelSizing for Rigidity (frequency)Sizing for StrengthSizing for Structural StabilityStructural Subsystem MassBackup Charts, ExtrasAnatomy of s/cSpacecraft Drawing in Launch VehicleDrawing of deployment phaseComposite Materials Composite Material PropertiesMass Estimation, Propulsion SubsystemMass Estimation, Propulsion SubsystemMass Estimation, Propulsion SubsystemDeployable Boom, equations for impact torqueStatic Beam DeflectionsRigid-body accelerationsCombining the loads to form quasi-static levelsLoads for a component, exampleSpinning upper stage example, contLoads for a component, cont