falcon 1 - Blackboard

USA
FALCON 1
1. IDENTIFICATION
1.1
Name
FALCON 1
1.2
Classification
¾ Family
¾ Series
¾ Version
:
:
:
FALCON
FALCON 1
FALCON 1
¾
¾
¾
Category :
Class
:
Type
:
SPACE LAUNCH VEHICLE
Small Launch Vehicle (SLV)
Expendable Launch Vehicle
1.3
Manufacturer
:
Space Exploration Technologies ( SpaceX)
1310 East Grand Avenue
EL SEGUNDO, CALIFORNIA 90245, USA
Telephone:
(310) 414-6555
Fax:
(310) 414-6552
1.4
Development manager :
Space Exploration Technologies ( SpaceX)
1310 East Grand Avenue
EL SEGUNDO, CALIFORNIA 90245, USA
Telephone: (310) 414-6555
Fax:
(310) 414-6552
1.5
Vehicle operator
:
Space Exploration Technologies ( SpaceX)
1310 East Grand Avenue
EL SEGUNDO, CALIFORNIA 90245, USA
Telephone: (310) 414-6555
Fax:
(310) 414-6552
1.6
Launch service agency :
Space Exploration Technologies ( SpaceX)
1310 East Grand Avenue
EL SEGUNDO, CALIFORNIA 90245, USA
Telephone: (310) 414-6555
Fax:
(310) 414-6552
1.7
Launch cost
6.7 M$ including launch range, third party insurance and standard
payload integration costs
:
2. STATUS
2.1
Vehicle status
:
Under development
2.2
Development period
:
2002-2005
2.3
First launch
:
Foreseen in 2005
December 2004
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3. PAYLOAD CAPABILITY AND CONSTRAINTS
3.1
Payload capability
3.1.1 Low Earth Orbits
Falcon is capable of delivering approximately 670 kg (1480 lb) payload into a reference low Earth orbit with
200 km circular altitude, when launched due East from Cape Canaveral.
With increasing altitude, a two-impulse insertion provides more payload capability by inserting into a lower,
slightly eccentric orbit and performing a circularization burn at apogee. The following graphs show
performance for both types of orbit insertion (Figure 1 to Figure 3) for a range of circular orbit altitudes.
ORBIT TYPE
LEO CIRCULAR
SSO CIRCULAR
Altitude
(km)
(Perigee/Apogee)
200
200
Inclination
28.5
-
Cap Canaveral
Vandenberg
670
570
(°)
Site
Payload mass (kg)
3.1.2 Geosynchronous Orbits
N/A
3.1.3 Injection accuracy
As a liquid propulsion vehicle, Falcon provides flexibility required for payload insertion into orbit with higher
eccentricity deploying multiple payloads into slightly different orbits. Target orbital insertion accuracy is
currently determined as follows:
¾
¾
¾
Inclination: +/- 0.1 deg
Perigee:
+/- 10 km
Apogee:
+/- 20 km
As for every new vehicle, insertion accuracy has to be verified during the first flights. The current estimates
are expected to improve since Falcon uses a GPS system for guidance correction.
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The following graphs show performance for both types of orbit insertion (Figure 1 and Figure 2) for a range
of circular orbit altitudes.
FIGURE 1 – PERFORMANCE FROM VANDENBERG
FIGURE 2 – PERFORMANCE FROM CAP CANAVERAL
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For orbit inclinations other than provided above, azimuth limitations have to be taken into account. The
azimuth limits depend on the type of payload and orbit. Figure 3 shows the payload performance as a
function of inclination and circular orbit altitude for typical azimuth limitations.
FIGURE 3 – CIRCULAR ORBIT PERFORMANCE FROM WESTERN RANGE AS A FUNCTION OF ORBIT
INCLINATION
3.2
Spacecraft orientation and separation
3.3
Payload interfaces
3.3.1 Payload compartments and adaptors
¾ Payload fairing description
Fairing dimensions
¾
¾
¾
¾
¾
Total length
Usable length
Primary diameter
Usable diameter
Mass
December 2004
:
:
:
:
:
3.5 m
2.8 m
1.5 m
1.3 m
-
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Dimensions are in meters and in inches inside the parentheses.
FIGURE 4 – PAYLOAD FAIRING DYNAMIC ENVELOPE
¾ Payload fairing separation
¾ Payload access provisions
3.4
Environments
3.4.1 Mechanical environment
During flight, payload will be submitted to a range of axial and lateral acceleration.
Axial acceleration is determined by the vehicle thrust history and drag, while max. lateral accelerations are
primarily determined by wind gusts, engine gimbal maneuvers, and other short-duration events.
Conservative loads used for payload design are summarized in Table 1 and axial acceleration vs. time is
plotted in Figure 5 for a nominal trajectory. Mission specific loads will be determined approximately four
months prior to launch with the coupled loads analysis (including specific trajectory details and payload mass
properties). These results should only be used to validate that the design loads envelope the mission specific
loads.
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Quasi-static Load Factors
Flight Event
Axial (g): Steady ± Dynamic (Total)
Lateral (g)
0.5
2.0
Lift Off
1.2 ± 0.4 (0.8 / 1.6)
0.50
Max qα
2.0 ± 1.0 (3.0 / 1.0)
0.75
Stage I burnout
6.4 ± 1.25 (7.7 / 5.2)
0.75
Stage II Ignition
3.2 ± .25 (3.0 / 3.4)* to 6.0 ± .25 (5.75 / 6.25)*
0.25
Stage II burnout
4.5 ± 0.5 (5.0 / 4.0)* to 6.5 ± 0.5 (7.0 / 6.0)*
0.25
Ground Handling
* Depending on payload mass and trajectory
TABLE 1 - SUMMARY OF PAYLOAD DESIGN C.G. LIMIT LOAD FACTORS
FIGURE 5 - NOMINAL STEADY STATE AXIAL ACCELERATION TIME HISTORY
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¾ Vibrations
The payload vibration environment is generated by acoustic noise in the fairing and also by engine and
aero-induced vibration that is transmitted through the vehicle structure. Because random vibration is very
difficult to predict analytically, preliminary Falcon environment was calculated using empirical correlations
coupled with scaling data from vehicles of similar size (physical and thrust), materials, and construction.
The generation and transmission of random vibrations scales closely with the acoustic power generation
of the engine, the materials comprising the skin, tanks, and fairing, and the type of joints used
respectively.
Based on this analysis, Falcon’s random vibration Maximum Predicted Environment (MPE) is shown
below in Figure 6. Note that these values include appropriate margins due to uncertainty and that this
data will be continuously refined as additional engine tests are performed. The corner frequencies and
slope values are summarized in Table 2.
Frequency (Hz)
PSD (g²/Hz)
Frequency Range (Hz)
PSD Slope
20
0.0043
0 – 20
0
100
0.020
20 - 100
+3 dB/oct
800
0.020
100 – 800
0
2000
0.0019
800 - 2000
-9 dB/oct
TABLE 2 - SUMMARY OF RANDOM VIBRATION PSD VALUES
FIGURE 6 - RANDOM VIBRATION SPECTRA
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3.4.2 Acoustic vibrations
During flight, the payload will be subjected to a varying acoustic environment. Levels are highest at lift off
and during transonic flight due to aerodynamic excitation. Falcon will make use of acoustic blanketing to
reduce the acoustic environment and a nominal (minimal) 2” thick blanket configuration is assumed for
predicted environment. Spectral energy methods were used to predict worst-case acoustic environment
below. A 1.5 dB margin is typically added for qualification and a 3 dB margin is typically added for testing.
A summary of acoustic MPE (Maximum Predicted Environment) is shown in Table 3 and plotted in Fig. 7.
1/3 Octave Frequency (Hz)
SPL (dB)
1/3 Octave Frequency (Hz)
SPL (dB)
20.0
25.0
31.5
40.0
50.0
63.0
80.0
100
125
160
200
250
315
400
104
107
110
113
115
117
119
120
121
122
122
122
121
119
500
630
800
1000
1250
1600
2000
2500
3150
4000
5000
6300
8000
10000
OASPL
117
115
115
112
107
104
102
101
100
101
101
100
100
100
131
TABLE 3 - SUMMARY OF PAYLOAD ACOUSTIC ENVIRONMENT ASSUMING NOMINAL 2” ACOUSTIC
BLANKETS
FIGURE 7 - SPL (SOUND PRESSURE LEVEL) SPECTRUM ASSUMING 2” ACOUSTIC BLANKETS
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3.4.3 Shock
There are four events during flight that are characterized as shock loads:
1) vehicle hold-down release at lift-off,
2) stage separation,
3) fairing separation, and
4) payload separation.
The baseline separation system for Falcon is the Lightband motorized ring system, which provides a very low
shock load (negligible compared pyrotechnic events). Data from Lightband is shown in Figure 8 (Shock
Response Spectrum (SRS)) below with a typical pyrotechnic shock spectrum shown for reference. Of the
other shock events, (1) and (2) are negligible for the payload relative to (3) due to the large distance and
number of joints over which shocks (1) and (2) will travel and dissipate. Max shock loading (3) is modeled
using by assuming an initial SRS given by the typical pyrotechnic curve in Figure 8 and scaling down via
standard correlations based on distance from source and joints. The resulting max. shock environment
predicted at payload interface is shown in Figure 9.
It should be noted that vibration test data of the Lightband system suggests very low response at the high
frequencies associated with shock events which will likely add considerable attenuation to any shock
acceleration loads. In addition, a damping and isolation system is being designed at the payload interface
which will further reduce any shock loads.
These systems are not modelled here and hence the SRS presented should be treated as very conservative.
FIGURE 8 AND 9- SHOCK RESPONSE AT SEPARATION PLANE DUE TO PAYLOAD SEPARATION
SYSTEM AND OTHER ORDNANCE EVENTS
3.4.4 Thermal environment
-
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3.4.5 Variation of static pressure under fairing
The fairing flight pressure profile is defined in Figure 10
FIGURE 10 - DEPRESSURIZATION ENVIRONMENT AND DEPRESSURIZATION RATES
3.5
Operation constraints
The standard launch integration process consists of the following:
¾ Contract signing and authority to proceed (Launch – 8 months or more)
-
Estimated Payload mass, volume, mission, operations and interface requirements
-
Safety information (Safety Program Plan; Design information: battery, ordnance, propellants; and
operations)
-
Mission analysis summary provided to the Customer within 30 days of contract
¾ Final payload design, including: mass, volume, structural characteristics, mission, operations, and
interface requirements (Launch – 6 months)
-
Payload to provide test verified structural dynamic model
¾ Payload readiness review for Range Safety (Launch – 4 months)
-
Launch site operations plan
-
Hazard analyses
¾ Verification (Launch – 3 months)
-
Review of Payload test data verifying compatibility with Falcon environments
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-
Coupled Payload and Falcon loads analysis completed
-
Confirm Payload interfaces as built are compatible with Falcon
-
Mission safety approval
¾ Pre-shipment review (Launch -1 month)
¾ Payload arrival at launch location (Launch – 2 weeks)
¾ Payload encapsulation (Launch – 3 days)
¾ Payload stacking (Launch – 2 days)
¾ Launch readiness Review (Launch – 1 day); Launch
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4. LAUNCH INFORMATION
4.1
Launch site
¾ Location
SpaceX has launch facilities at Space Launch Complex 3W at Vandenberg Air Force base for high
inclination missions, Launch Complex 46 at Cape Canaveral Air Force Station for mid inclination
missions, and the Reagan Test Site on the Kwajalein Atoll for very low inclination missions.
FIGURE 11 – KWAJALEIN LAUNCH SITE LOCATION
¾ Payload Processing Facility
The Space Exploration Technologies Payload processing facility will be made available for non
hazardous Payload operations for up to 3 weeks prior to launch from each of our sites. If additional time is
needed, then the Payload must procure a processing facility independently.
The payload processing room at SLC 3W consists of a high bay 60 ft long, 20 ft wide, 20 ft high.
Personnel access doors provide access to the processing room from the launch vehicle/payload checkout
control room. The outside cargo entrance door is 16 ft wide by 16 ft high which leads into a 20 ft wide, 20
ft long, 20 ft high anti-room. A bridge crane with an 18 ft hook height and 5 ton capacity is available for
handling spacecraft and associated equipment. The processing room provides a class 100,000 laminar
flow cleanroom 40 ft long, 20 ft wide, 20 ft high. Temperature of 60° - 80° F ± 2° (controlled) and humidity
30% - 60% (monitored only) with a differential pressure of .05 inches of water minimum with all doors and
pass-throughs closed.
The Payload processing facilities will be consists of a common work area (anti-room) and the clean room
bay. The common work area of the processing facility is dedicated for spacecraft ground support
equipment unloading, unpacking/packing and intermediate storage of empty cargo container. The working
area shall be equipped with a crane to lift the Payload onto the adapter.
SpaceX will monitor relative humidity, temperature and cleanliness in the payload processing facility. This
is true with the exception of periods when the satellite is mated to the launcher second stage. Monitoring
data will be made available to the payload customer. In order to ensure this, satellite preparation will be
performed using clean processes.
The processing facility has the following characteristics:
-
Smooth, continuous floor surface (anti-static)
-
Illumination equals 300 Lux (with localized capability up to 1000 Lux)
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FALCON 1
-
Uninterrupted power supply for satellite control and test equipment
-
Fire protection
-
Emergency exit and illumination
-
Controlled access and security
Sequence of flight events
FLIGHT Time
Event
0 s Stage 1 Ignition and Liftoff
169 s Stage 1 Burnout, Stage Separation
174 s Stage 2 Ignition
194 s Fairing Separation
552 s Stage 2 Burnout
570 s Payload Deployment
1114 s Stage 1 Splshdown
FIGURE 11 - FALCON SAMPLE FLIGHT PROFILE – DIRECT INJECTION
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4.3
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Launch record data
¾ Failures
:-
¾ Provisional reliability : ¾ Success ratio
4.4
:-
Planned launches
December 2004
LAUNCH DATE
VEHICLE
DEPARTURE POINT
3rd quarter of 2005
Falcon 1
Vandenberg
3rd quarter of 2005
Falcon 1
Marshall Islands
4th quarter of 2005
Falcon 1
Marshall Islands
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5. DESCRIPTION
5.1
Launch vehicle
FIGURE 12 – FALCON LAUNCH VEHICLE
5.2
¾
¾
¾
Overall vehicle
Overall length
Maximum diameter
Lift-off mass
December 2004
: 21.3 m
: 1.7 m
: 27.2 t
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5.3
FALCON 1
General characteristics of the stages
STAGE
1
2
-
-
SpaceX
SpaceX
-
-
Diameter (m)
1.7
1.7
Dry mass (kg)
-
-
Liquid
Liquid
Designation
Manufacturer
Length (m)
Propellant:
¾ Type
¾ Fuel
¾ Oxidizer
Kerosene (RP-1) Kerosene (RP-1)
Oxygene
Oxygene
¾ Fuel
-
-
¾ Oxidizer
-
-
TOTAL
-
-
-
-
Propellant mass (kg)
Tank pressure (bar)
Total lift-off mass (kg)
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5.4
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Propulsion
STAGE
1
2
Merlin
Kestrel
SpaceX
SpaceX
Number of engines
1
1
Engine mass (kg)
-
-
Turbopump
Tank pressure
Mixture ratio
-
-
Chamber pressure (bar)
-
-
Cooling
-
Ablative
¾ Sea level
255
-
¾ Vacuum
304
327
¾ Sea level
342
-
¾ Vacuum
409
31
Burning time (s)
169
378
-
-
No
Yes
Designation
Manufacturer
Feed syst. type
Specific impulse (s)
Thrust (kN)
Nozzle expansion ratio
Restart capability
5.5
Guidance and control
5.5.1 Guidance
5.5.2 Control
The second stage attitude and rate accuracies at separation are:
¾
¾
¾
Roll
: +/- 2 degrees
Pitch/Yaw : +/- 0.5 degrees
Body rates : +/- 0.1 deg/sec/axis
STAGE
Pitch, yaw
Roll
December 2004
1
2
-
-
By gimballing the
nozzle
-
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6. DATA SOURCE REFERENCES
1
-
Falcon Launch Vehicle - Payload User’s Guide - Rev 2 October 2004
2
-
http://www.spacex.com
December 2004
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