USA FALCON 1. Fax: (310) Telephone: (310) Fax: (310) Telephone: (310) Fax: (310)

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1 1. IDENTIFICATION 1.1 Name FALCON Classification Family : FALCON Series : FALCON 1 Version : FALCON 1 Category : SPACE LAUNCH VEHICLE Class : Small Launch Vehicle (SLV) Type : Expendable Launch Vehicle 1.3 Manufacturer : Space Exploration Technologies ( SpaceX) 1310 East Grand Avenue EL SEGUNDO, CALIFORNIA 90245, USA Telephone: (310) Fax: (310) Development manager : Space Exploration Technologies ( SpaceX) 1310 East Grand Avenue EL SEGUNDO, CALIFORNIA 90245, USA Telephone: (310) Fax: (310) Vehicle operator : Space Exploration Technologies ( SpaceX) 1310 East Grand Avenue EL SEGUNDO, CALIFORNIA 90245, USA Telephone: (310) Fax: (310) Launch service agency : Space Exploration Technologies ( SpaceX) 1310 East Grand Avenue EL SEGUNDO, CALIFORNIA 90245, USA Telephone: (310) Fax: (310) 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 : First launch : Foreseen in 2005 December 2004 Page 1

2 3. PAYLOAD CAPABILITY AND CONSTRAINTS 3.1 Payload capability 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) Inclination ( ) Site Payload mass (kg) Cap Canaveral Vandenberg Geosynchronous Orbits N/A 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: 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. December 2004 Page 2

3 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 December 2004 Page 3

4 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 Payload interfaces Payload compartments and adaptors Payload fairing description Fairing dimensions Total length : 3.5 m Usable length : 2.8 m Primary diameter : 1.5 m Usable diameter : 1.3 m Mass : - December 2004 Page 4

5 Dimensions are in meters and in inches inside the parentheses. FIGURE 4 PAYLOAD FAIRING DYNAMIC ENVELOPE Payload fairing separation - Payload access provisions Environments 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. December 2004 Page 5

6 Flight Event Quasi-static Load Factors Axial (g): Steady ± Dynamic (Total) Lateral (g) Ground Handling Lift Off 1.2 ± 0.4 (0.8 / 1.6) 0.50 Max q α 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 * 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 December 2004 Page 6

7 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 db/oct db/oct TABLE 2 - SUMMARY OF RANDOM VIBRATION PSD VALUES FIGURE 6 - RANDOM VIBRATION SPECTRA December 2004 Page 7

8 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) OASPL 131 TABLE 3 - SUMMARY OF PAYLOAD ACOUSTIC ENVIRONMENT ASSUMING NOMINAL 2 ACOUSTIC BLANKETS FIGURE 7 - SPL (SOUND PRESSURE LEVEL) SPECTRUM ASSUMING 2 ACOUSTIC BLANKETS December 2004 Page 8

9 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 Thermal environment - December 2004 Page 9

10 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: 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 Final payload design, including: mass, volume, structural characteristics, mission, operations, and interface requirements (Launch 6 months) - Payload to provide test verified structural dynamic model Payload readiness review for Range Safety (Launch 4 months) - Launch site operations plan - Hazard analyses Verification (Launch 3 months) - Review of Payload test data verifying compatibility with Falcon environments December 2004 Page 10

11 - Coupled Payload and Falcon loads analysis completed - Confirm Payload interfaces as built are compatible with Falcon - Mission safety approval Pre-shipment review (Launch -1 month) Payload arrival at launch location (Launch 2 weeks) Payload encapsulation (Launch 3 days) Payload stacking (Launch 2 days) Launch readiness Review (Launch 1 day); Launch December 2004 Page 11

12 4. LAUNCH INFORMATION 4.1 Launch site 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. Payload Processing Facility FIGURE 11 KWAJALEIN LAUNCH SITE LOCATION 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 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) December 2004 Page 12

13 - Uninterrupted power supply for satellite control and test equipment - Fire protection - Emergency exit and illumination - Controlled access and security 4.2 Sequence of flight events FLIGHT Time 0 s 169 s 174 s 194 s 552 s 570 s 1114 s Event Stage 1 Ignition and Liftoff Stage 1 Burnout, Stage Separation Stage 2 Ignition Fairing Separation Stage 2 Burnout Payload Deployment Stage 1 Splshdown FIGURE 11 - FALCON SAMPLE FLIGHT PROFILE DIRECT INJECTION December 2004 Page 13

14 4.3 Launch record data - Failures : - Provisional reliability : - Success ratio : Planned launches LAUNCH DATE VEHICLE DEPARTURE POINT 3 rd quarter of rd quarter of th quarter of 2005 Falcon 1 Falcon 1 Falcon 1 Vandenberg Marshall Islands Marshall Islands December 2004 Page 14

15 5. DESCRIPTION 5.1 Launch vehicle FIGURE 12 FALCON LAUNCH VEHICLE 5.2 Overall vehicle Overall length Maximum diameter Lift-off mass : 21.3 m : 1.7 m : 27.2 t December 2004 Page 15

16 5.3 General characteristics of the stages STAGE 1 2 Designation - - Manufacturer SpaceX SpaceX Length (m) - - Diameter (m) Dry mass (kg) - - Propellant: Type Liquid Liquid Fuel Kerosene (RP-1) Kerosene (RP-1) Oxidizer Oxygene Oxygene Propellant mass (kg) Fuel - - Oxidizer - - TOTAL - - Tank pressure (bar) Total lift-off mass (kg) - - December 2004 Page 16

17 5.4 Propulsion STAGE 1 2 Designation Merlin Kestrel Manufacturer SpaceX SpaceX Number of engines 1 1 Engine mass (kg) - - Feed syst. type Turbopump Tank pressure Mixture ratio - - Chamber pressure (bar) - - Cooling - Ablative Specific impulse (s) Sea level Vacuum Thrust (kn) Sea level Vacuum Burning time (s) Nozzle expansion ratio - - Restart capability No Yes 5.5 Guidance and control Guidance Control The second stage attitude and rate accuracies at separation are: Roll : +/- 2 degrees Pitch/Yaw : +/- 0.5 degrees Body rates : +/- 0.1 deg/sec/axis STAGE 1 2 Pitch, yaw - - Roll By gimballing the nozzle - December 2004 Page 17

18 6. DATA SOURCE REFERENCES 1 - Falcon Launch Vehicle - Payload User s Guide - Rev 2 October December 2004 Page 18