1. IDENTIFICATION 1.1 Name DELTA 2-6925 1.2 Classification Family : DELTA Series : DELTA 2 Version : 6925 Category : SPACE LAUNCH VEHICLE Class : Medium Launch Vehicle (MLV) Type : Expendable Launch Vehicle (ELV) 1.3 Manufacturer : Mc DONNELL DOUGLAS SPACE SYSTEMS Company (MDSSC) 5301 Bolsa Avenue HUNTINGTON BEACH CALIFORNIA 92647 1.4 Development manager : U.S. Air Force Space Division L.A. Air Force Station, P.O. Box 92960 LOS ANGELES, CALIFORNIA 90009-2960 1.5 Vehicle operator : Mc DONNELL DOUGLAS SPACE SYSTEMS Company (with USAF or with NASA) 1.6 Launch service agency : Mc DONNELL DOUGLAS SPACE SYSTEMS Company 5301 Bolsa Avenue HUNTINGTON BEACH CALIFORNIA 92647 Tel. : (714) 896-3311 Telex : 67-8426 Fax : 714-896 - 1315 1.7 Launch cost : Between 45 and 50 M$ (1987) on the basis of one launch of a DELTA 2 serie launcher every 60 days. December 1992 Page 1
2. STATUS 2.1 Vehicle status : Out of service 2.2 Development period : Started January 1987 2.3 First launch : 15.02.1989 (success) 3. PAYLOAD CAPABILITY AND CONSTRAINTS 3.1 Payload capability 3.1.1 Low Earth Orbits - 3.1.2 Geosynchronous and Interplanetary Orbits GEOSYNCHRONOUS TRANSFER ORBIT AND PLANETERY CAPABILITIES FROM ESMC FIGURE 1 FIGURE 2 PERIGEE VELOCITY CAPABILITY FROM ESMC AND WSMC RESPECTIVELY FIGURE 3 FIGURE 4 December 1992 Page 2
APOGEE ALTITUDE CAPABILITY FROM ESMC AND WSMC RESPECTIVELY 3.1.3 Injection accuary FIGURE 5 FIGURE 6 All DELTA II configurations, including DELTA 2-6925, employ the DELTA Inertial Guidance System (DIGS) mounted in the second-stage guidance compartment. Orbit accuary for free-stage missions is primarily affected by the pointing and impulse error resulting from the upper stage solide motor burn. To assist the user in mission planning, 3σ apogee altitude dispersion data for a geosynchronous transfer orbit are presented in the figure hereunder, as a function of transfer orbit inclination. The band shown in the figure indicates a typical range of pointing error at upper stage ignition. The pointing error for any given mission depends on upper stage/spacecraft mass properties and spin rates. Past DELTA experience, however, shows that most missions will typically fall within a 1 to 2.5 degrees pointing error band. Transfer orbit perigee altitude errors for geosynchronous missions are typically < 5.6 km. Three-sigma transfer orbit inclination errors over the inclination range shown in the figure are typically 0.2 < i < 0.6 degrees. These data should be used as general accuracy indicators only. Detailed analyses are performed for each specific mission, including the effects of individual mission requirements to define more precisely the accuracy to be expected. Users are invited to contact the II Program Office for further information on orbit accuracy. FIGURE 7 - THREE-STAGE TYPICAL SYNCHRONOUS APOGEE ALTITUDE DEVIATION (km) - ESMC December 1989 Page 3
3.2 Spacecraft orientation and separation The third-stage DELTA separation sequencing system consists of a pyrotechnic timer to effect spacecraft separation for redundancy, two timers are located on the attach fitting and incorporated into the third-stage sequencing system. The sequencing system is initiated by a signal from the guidance computer before second / third-stage separation. FIGURE 8 - DELTA 2-6925 AND DELTA 2-7925 SEPARATION SYSTEMS Relative payload separation velocity: 0,61 to 2,4 m/s Deployement mechanism type: 4 springs release 3.3 Payload interfaces 3.3.1 Payload compartments and adaptors Payload fairing description: the fairing is an all-aluminium structure fabricated in two half-shells, consisting of a hemispherical nose cap with a biconic section. Dimensions and mass of the fairing : - length: 8 488 mm - diameter: from 2 438 mm at the base to 2 896 mm at the midest portion - mass: 839 kg including thermal and acoustic insulation December 1989 Page 4
The allowable spacecraft envelope within the confines of the fairing is shown herebelow: FIGURE 9 - PAYLOAD ENVELOPE, PAM-D STAR 48B CONFIGURATION (3712 ATTACH FITTING) The Payload Attach Fitting (PAF) 3712 is the interface between the upper stage motor of the three-stage versions of DELTA II and the spacecraft. The 3712 fitting is available with three forward flange configurations designated 3712 A, 3712 B, 3712 C. The figure hereunder shows the capabilities of the three configurations in terms of spacecraft weight, cg location and the following maximum clamp assembly preload associated with each configuration: 3712 A 3 084 kg 3712 B 3 855 kg 3712 C 1 272 kg December 1989 Page 5
FIGURE 10 - CAPABILITIES OF THE THREE CONFIGURATIONS The actual selection of attach fitting configuration will be made after discussions with the spacecraft contractor. As an option, the PAF is available without an NCS (Nutation Control System). New payload attach fittings are currently being studied; they take into account the use of the separation clamp assembly interfaces that have been qualified for STS. They are listed below: APPROXIMATE DIAMETER (MM) MAXIMUM FLIGHT PRELOAD (KG) SPACECRAFT PAF FLANGE ANGLE (DEG) 115 3084.5 15 123 2585.5 20 136 3538 20 146 2993.7 20 164 3538 15 A variety of removable doors can be provided in the fairing to permit limited access to the spacecraft following fairing installation. Four standard access doors exit in the baseline fairing configuration.the RF windows and access doors do note contain acoustic or thermal blanket material. Additionnal access doors can be envisioned. December 1989 Page 6
3.4 Environments 3.4.1 Mechanical environment Steady - state acceleration For a three-stage DELTA vehicle the maximum steady-state acceleration occures at the end of thirdstage flight for payloads up to 1 180 kg. Above this weight, the maximum acceleration occurs at the end of the first-stage burn. FIGURE 11 - AXIAL STEADY-STAGE ACCELERATION AT UPPER STAGE BURNOUT Vibration The maximum expected sinusoïdal vibration flight levels (3 σ) are as follows: FIGURE 12 - AXIAL STEADY-STAGE ACCELERATION AT MECO VERSUS SECOND- STAGE PAYLOAD WEIGHT FREQUENCY (Hz) ACCELERATION (g's) (zero to peak) Thrust axis 5 to 6.2 6.2 to 100 0.5 IN (Double amplitude) 1.0 Lateral axis 5 to 100 0.7 The spacecraft random vibration is best simulated by a properly performed acoustic test sed on the environment described hereafter. Specific random vibration interface levels may be requested. 3.4.2 Acoustic vibrations The acoustic environment for the 6925 version of the DELTA 2 vehicle is shown hereunder. FIGURE 13 - INTERNAL FAIRING ACOUSTIC ENVIRONMENT FOR DELTA 2-6925 December 1989 Page 7
3.4.3 Shock environment FIGURE 14 - MAXIMUM FLIGHT SHOCK LEVELS FOR THE DELTA 2-6925 AND DELTA 2-7925 VEHICLES DUE TO CLAMPBAND SEPARATION SYSTEM 3.4.4 Thermal environment Fairing FIGURE 15 - VEHICLE FAIRING INTERNAL WALL TEMPERATURE AND EMITTANCE December 1989 Page 8
Upper-stage motor and spin-rockets The motor Plume subjects the spacecraft to a heat flux profile depicted in the following figures. FIGURE 16 - PAM-D PLUME RADIATION AT THE SPACECRAFT SEPARATION PLANE VERSUS BURN TIME FIGURE 17 - PAM-D PLUME RADIATION AT THE SPACECRAFT SEPARATION PLANE VERSUS CENTERLINE DISTANCE The spin rocket plumes subject the spacecraft to a maximum heat of 2 837 W/m² at the separation plane. This heat flux is a 1 s duration pulse. The upper-stage motor case temperatures are detailed herebelow. FIGURE 18 - UPPER STAGE MOTOR CASE TEMPERATURES 3.4.5 Variation of static pressure within the fairing Venting of the fairing is provided by a 64.5 cm² vent opening in the interstage and by other leak paths in the vehicle. December 1989 Page 9
The resulting fairing internal pressure history in represented in the following figure: FIGURE 19 - DELTA 2 9.5 ft. FAIRING INTERNAL PRESSURE LIMITS 3.4.6 Spacecraft compatibility tests Sinusoïdal vibration test levels Acoustic test level December 1989 Page 10
3.5 Operation constraints Launch rate capability (including all DELTA II versions): 12 in 1990, up to 18 in 1991. Ground constraints The spacecraft agency must provide its own test equipment for spacecraft preparations, including telemetry and ground stations. Integration process FIGURE 20 - INTEGRATION PROCESS A typical launch site operations flow involves about 73 working days. The spacecraft is not required for mate with the DELTA third stage until about 10 working days prior to launch. FIGURE 21 - TYPICAL LAUNCH OPERATION FLOW December 1989 Page 11
4. LAUNCH INFORMATION 4.1 Launch site Dual pad launch capabilities are available in Florida at Launch Complex 17 on the Cape Canaveral Air Force Station (CCAFS). In addition, a single launch pad is available at Vandenberg Air Force Base (VAFB) in California to accomodate polar or other high inclination launches. Cape Canaveral Air Force Station, a part of the ESMC (Eastern Space and Missile Center), is located approximately 80 km east of Orlando (Florida). Launch Complex 17 consists of two launch pads (17 A and 17 B), a blockhouse, ready room, shops, and other facilities needed to prepare, service and launch the DELTA vehicle. Launch information is shown hereafter : LAUNCH PAD LATITUDE (degrees North) LONGITUDE (degrees West) LAUNCHER AZIMUT 17 A 28.44687 80.56516 115.2 17 B 28.44559 80.56605 115.0 Launch Complex 17 is convenient for orbit inclinations from 28.5 to 51. In addition to the facilities required for the DELTA II launch vehicle, specialized facilities are provided for the spacecraft: Payload processing facilities: - NASA - Provided hangars AO, AM, or AE, - ASTROTECH Space Operations in Titus-ville, Hazardous processing facilities: - NASA - Provided payload spin test facilities: Explosive Safe Area 60 (ESA 60), Spacecraft Assembly and Encapsulation Facility 2 (SAEF 2), Cargo Handling Storage Facility at KSC or CCAFS; - ASTROTECH Space Operations ASTROTECH of these facilities is controlled by the respective owners. December 1989 Page 12
FIGURE 22 - LAUNCH COMPLEX 17, CCAFS FIGURE 23 - DELTA SPACECRAFT CHECKOUT FACILITIES December 1989 Page 13
Vandenberg Air Force Base (CALIFORNIA) The pad available at VAFB for DELTA launches, the Space Launch Complex SLC-2, is convenient for orbit inclinations from 63 to 145. LAUNCH PAD LATITUDE (degrees North) LONGITUDE (degrees West) LAUNCH AZIMUT SLC-2 34.75 120.62 259.5 FIGURE 24 - VANDENBERG AIR FORCE BASE FIGURE 25 - SPACE LAUNCH COMPLEX SLC-2 WSMC - PLAN VIEW December 1989 Page 14
FIGURE 26 - SPACECRAFT SUPPORT AREA 4.2 Sequence of flight events Typical sequence of events for a DELTA 2-6925 mission launched from ESMCR. FLIGHT TIME EVENTS To Main engine ignition Solid motor ignition (6 solids) 56 Solid motor burnout (6 solids) 62 Solid motor ignition (3 solids) 63/64 Solid motor separation (3/3 solids) 118 Solid motor burnout (3 solids) 124 Solid motor separation (3 solids) 265 Main engine cutoff 273 Blow stage 1/2 separation bolts 278 2nd stage ignition 305 Fairing separation 692 2nd stage engine cutoff 1 300 2nd stage engine restart 1 310 2nd stage engine cutoff command 1 360 Fire spin rockets, start 3rd stage sequencer 1 362 3rd stage separation 1 400 3rd stage burnout 1 600 Spacecraft separation December 1992 Page 15
4.3 Launch record data LAUNCH DATE NUMBER OF SATELLITES ORBIT RESULT REMARK Failures: none 14.02.89 1 Circular Success 10.06.89 1 Circular Success 18.08.89 1 Circular Success 21.10.89 1 Circular Success 11.12.89 1 Circular Success 24.01.90 1 Circular Success 14.02.90 (1) 1 LEO Success DELTA 2 6920 26.03.90 1 Circular Success 13.04.90 1 GEO Success 01.06.90 (1) 1 LEO Success DELTA 2 6920 02.08.90 1 Circular Success 18.08.90 1 GEO Success 01.10.90 1 Circular Success 30.10.90 1 Circular Success 08.03.91 (2) 1 GEO Success PAM-D2 07.06.92 (1) 1 LEO Success DELTA 2 6920 24.07.92 1 GTO Success (1) 2 stage vehicle (DELTA 2-6920) (2) 3rd stage: PAM-D2 LAUNCH DATE RESULT CAUSE 24.01.94. Success ratio: 100% (on 17 launches) 4.4 Planned launches None of the 6 900 series. December 1992 Page 16
5. DESCRIPTION 5.1 Launch vehicle FIGURE 27 - VIEW OF DELTA 2-6925 5.2 Overall vehicle Overall length Maximum diameter Lift-off mass (approx.) : 39.62 m : 2.44 m : 220 t December 1992 Page 17
5.3 General characteristics of the stages STAGE 0 1 2 3 Designation CASTOR 4 A - - PAM-D Manufacturer THIOKOL MDSSC MDSSC MDSSC Length (m) 10.97 26.08 5.96 Diameter (m) 1.02 2.44 2.44 1.20 Dry mass (t) 1.5 Propellant : Type Solid Liquid Liquid Solid TP-H8299 (Storable) (Storable) Fuel CTPB RP-1 UDMH (A-50) Oxidizer Oxygen N 2 O 4 HTPB 2000 Propellant mass (t) : Total 10 96 5.9 Max.2 Fuel Oxidizer Water Tank pressure (bar) Total lift-off mass (t) 11.5 6.9 2.2 Upper part DESIGNATION SPIN TABLE PAYLOAD FAIRING PAYLOAD ATTACH FITTING Manufacturer Mass 839 kg Launch vehicle growth: yes, there are 2 improved versions of DELTA 2-6925 planned in the future: DELTA 2-7925 and DELTA 2A. December 1989 Page 18
5.4 Propulsion STAGE 0 1 2 3 Designation TX - 780 RS - 27 AJ 10-118 K STAR 48 Manufacturer THIOKOL Rocketdyne AEROJET THIOKOL Number of engines 9 1 1 1 Engine mass (kg) Feed syst.type - - Mixture ratio - Chamber pressure (bar) 47.6 Cooling Specific impulse (s): sea level 230 262 vacuum 303.5 319 Thrust (kn): sea level 480 919 vacuum 378 912 42 66.7 Burning time (s) 52.5 228 450 85.3 Nozzle expansion ratio Restart capability 8.2 8:1 140 53.9 Yes 5.5 Guidance and control 5.5.1 Guidance Inertial guidance unit on first and second stages + DELCO processor unit. 5.5.2 Control STAGE 0 1 2 3 Pitch, yaw TVC Gimballing Gimballing Gimballing Roll - - - Spinstabilized 30 to 10 rpm Precision - - - - December 1989 Page 19
6. DATA SOURCE REFERENCES 1 - DELTA 2 spacecraft users manual - July 1987 2 - AIAA 89-2422: Castor solid rocket motors for launch vehicle propulsion - July 10-12, 1989 3 - DELTA 2: A new generation - Mc DONNELL DOUGLAS SPACE SYSTEMS Company brochure - March 1989 4 - DELTA 2: Brochure released by Mc DONNELL DOUGLAS ASTRONAUTICS Company - June 1988 December 1989 Page 20