DELTA DELTA , 7326, 7425, H, 7926, 7926H

Transcription

1 . IDENTIFICATION. Name : DELTA DELTA 27325, 7326, 7425, H, 7926, 7926H.2 Classification Family : DELTA Series : DELTA 2 Versions : DELTA 27abc(H) Category : SPACE LAUNCH VEHICLE Class : Medium Launch Vehicle (MLV) Type : Expendable Launch Vehicle (ELV) a : Quantity of solid rocket motors b : Type of second stage : 2 = Aerojet AJ08K engine c : Type of third stage : 5 = Star48B ; 6 = Star37 FM ; H = Heavy configuration with GEM46 SRM ; Dash nr Fairing designator.3 Manufacturer : The Boeing Company 530 Bolsa Avenue Huntington Beach CA Telephone : (74) Fax : (74) Development manager : The Boeing Company.5 Vehicle operator : The Boeing Company.6 Launch service agency : Delta Launch Services c/o The Boeing Company 530 Bolsa Avenue, MC H04C426 Huntington Beach, CA deltalaunchservices@boeing.com.7 Launch costs : M$ 2. STATUS 2. Vehicle status : Operational 2.2 Development period : Started January First launch : December 2003 Page

2 3. PAYLOAD CAPABILITY AND CONSTRAINTS 3. Payload capability 3.. Low Earth Orbits 3..2 Geosynchronous and Interplanetary Orbits Vehicle designation Third stage Fairing GTO 85 x km CCAFS () i = 28.7 Interplanetary C3 = 0.4 km2/sec2 CCAFS i = 28.7 Molniya orbit 370 x km VAFB (2) i = Star48B 2.9m 3.0m Star37FM 2.9m 3.0m N/A (3) N/A (3) N/A N/A Star48B 2.9m 3.0m Star37FM 2.9m 3.0m N/A N/A L Star48B 2.9m 3.0m 3.0Lm L Star37FM 2.9m 3.0m 3.0Lm H 7925H0 7925H0L 7926H 7926H0 7926H0L Star48B 2.9m 3.0m 3.0Lm Star37FM 2.9m 3.0m 3.0Lm Currently not available from VAFB TABLE : THREESTAGE MISSION CAPABILITIES () CCAFS : Cape Canaveral Air Force Station. (2) VAFB : Vandenberg Air Force Base. (3) Not available, exceeds maximum allowable Star48B motor offload capability. December 2003 Page 2

3 Next figures show the three stages Delta 2 capabilities for launches from Eastern Range (ER) site. FIGURE : DELTA 2732X/742XH GTO INCLINATION CAPABILITY ER December 2003 Page 3

4 FIGURE 2 : DELTA 2792X/792XH VEHICLE, GTO INCLINATION CAPABILITY ER 3..3 Injection accuracy All Delta 2 configurations employ the RIFCA mounted in the secondstage guidance compartment. This system provides precise pointing and orbit accuracy for both two and three stage missions. The typical pointing error at thirdstage ignition is approximately.5 for the Star48B and 2.0 for the Star37FM motor. Deviations from nominal apogee altitude using the 7300, 7400, 7900, and 7900H launch vehicles for GTO mission from eastern range launch site are shown in Figure 3. The transfer orbit inclination error is typically from ±0.2 to ± 0.6 over the range shown, while the perigee altitude variation is typically about ± 9.3 km. All errors are 3σ values. December 2003 Page 4

5 FIGURE 3 : GTO DEVIATIONS CAPABILITY ER 3.2 Spacecraft orientation and separation Thermal control manoeuvres : yes Nominal payload separation velocity : ~0.3 m/s Rotation rate : 4 rpm Deployment mechanism type : explosive nuts or by release of a Vband clamp December 2003 Page 5

6 3.3 Payload interfaces 3.3. Payload compartments and adaptors Payload Attach Fittings (PAF) There are four standard PAFs available for threestage missions. The 372 PAF comes in three forward flange configurations, designated 372A, 372B, and 372C. The 3724 PAF is available with one forward flange configuration, designated 3724C. The first two digits of each PAF s designation indicate its payload interface diameter in inches, and the last two digits indicate the PAF s height in inches. All PAFs are designed such that payload electrical interfaces and separation springs can be located to accommodate specific customer requirements. Model All dimensions are in mm/(in.) Separation mechanism Features FIGURE 4 : DELTA 2 PAYLOAD ADAPTORS AND INTERFACES The maximum clampband flight preload for the 372 and 3724 configurations is given in table 2 : PAF Max flight preload (N) Spacecraft PAF flange angle 372A 30, B / 372C / 3724C 7, TABLE 2 : MAXIMUM CLAMPBAND ASSEMBLY PRELOAD December 2003 Page 6

7 Payload fairings The Delta 2 offers three fairings : a 2.9 m diameter metallic fairing (bisector) and a 3.0 m diameter composite fairing that comes in 2 different lengths. The stretched 3 m composite fairing designated 0L has a reshaped nose cone and a cylindrical section 0.9 m longer than the standard 3 m version. The fairings incorporate interior acoustic absorption blankets as well as flightproven contaminationfree separation joints. The Boeing Company supplies missionspecific modifications to the fairings as required by the customer. These include access doors, additional acoustic blankets, and RF windows. FIGURE 5 : DELTA 2 PAYLOAD FAIRINGS December 2003 Page 7

8 The 2.9 m diameter payload fairing The 2.9 m diameter fairing is an aluminium skin and stringer structure fabricated in two halfshells. These shells consist of a hemispherical nose cap, a biconic section, a cylindrical 2896 mm diameter center section, a 30 deg conical transition and a cylindrical base section having the 2438 mm core vehicle diameter. The biconic section is a ringstiffened monocoque structure; onehalf of which is fiberglass covered with a removable aluminium foil lining to create a RF window. The cylindrical base section is an integrally stiffened isogrid structure, and the cylindrical center section has a skin and stringer construction. The fairing has an overall length of 8488 mm. The halfshells are joined by a contaminationfree linear piston/cylinder thrusting separation system that runs longitudinally the full length of the fairing. Two functionally redundant explosive bolt assemblies provide structural continuity at the fairing base ring. Four functionally redundant explosive bolt assemblies (two each) provide circumferential structural continuity at the 30 deg transition section between the 2896 mm Φ section and the 2438 mm Φ section. The fairing halfshells are jettisoned by actuation of the base and transition separation nuts and by the detonating fuse in the thrusting joint cylinder rail cavity. Two 457 mm by 457 mm access doors for secondstage access are part of the baseline fairing configuration. FIGURE 6 : 2.9 m DIAMETER PAYLOAD FAIRING December 2003 Page 8

9 FIGURE 7 : PAYLOAD STATIC ENVELOPE, 2.9m Φ FAIRING (372 PAF) December 2003 Page 9

10 The 3 m diameter payload fairing The 3 m diameter fairing is a composite sandwich structure that separates into bisectors. Each bisector is constructed in a single cocured layup, eliminating the need for moduletomodule manufacturing joints and intermediate ring stiffeners. The resulting smooth inside skin enables the flexibility to install missionunique access doors almost anywhere in the cylindrical portion of the fairing. An RF window can be accommodated, similar to missionunique access doors. The bisectors are joined by a contaminationfree linear piston/cylinder thrusting separation system that runs longitudinally the full length of the fairing. Two functionally redundant explosive bolt assemblies provide the structural continuity at the fairing base ring. The fairing bisectors are jettisoned by actuation of the base separation nuts, and by the detonating fuse in the thrusting joint cylinder rail cavity. Two standard 457 mm diameter access doors are part of the baseline fairing configuration for secondstage access. Additional 60 mm Φ doors can be provided in the fairing cylindrical section for spacecraft access after encapsulation. FIGURE 8 : PAYLOAD STATIC ENVELOPE, 3 m Φ FAIRING (372 PAF) December 2003 Page 0

11 The stretched 3 m diameter payload fairing 0L The stretched 3 m Φ fairing, designated 0L, is available for payloads requiring a longer envelope than the 3 m Φ fairing. The 0L fairing is also a composite sandwich structure that separates into bisectors. The cylindrical section is lengthened by m, making the overall length 0.36 m longer than the 3m Φ fairing. The dualpayload attach fitting (DPAF) is also available for the stretched 3 m Φ (0L) fairing. The allowable static spacecraft envelopes are shown in Figure 9 : FIGURE 9 : PAYLOAD STATIC ENVELOPE, 3 m Φ STRETCHED COMPOSITE FAIRING (0L) (372 PAF) December 2003 Page

12 3.4 Environments 3.4. Mechanical environment SteadyState acceleration The maximum steadystate acceleration occurs at the end of thirdstage flight for payloads up to kg for the Star48B and 60.0 kg for the Star37FM. Above this weight, the maximum acceleration occurs at MECO. Steadystate axial acceleration versus payload weight at thirdstage motor burnout is shown in Figure 0. FIGURE 0 : AXIAL STEADYSTATE ACCELERATION AT THIRDSTAGE BURNOUT Combined loads Dynamic excitations, which occur predominantly during liftoff and transonic periods of Delta flight, are superimposed on steadystate accelerations to produce combined accelerations. To avoid dynamic coupling between lowfrequency vehicle and spacecraft modes, the stiffness of the spacecraft structure should have fundamental frequencies above 35 Hz in the thrust axis and 2 Hz for a twostage spacecraft in the lateral axis for a spacecraft hardmounted at the spacecraft separation plane (without PAF and separation clamp). In addition, secondary structure mode frequencies should be above 35 Hz. AXIS LIFTOFF / TRANSONIC Lateral ± 2.0 / ± 2.5 () Thrust ± 2.5 / 0.2 (2) () Lateral load factor to provide correct bending moment at spacecraft separation plane (2) Plus indicates compression load and minus indicates tension load December 2003 Page 2

13 3.4.2 Acoustic vibrations The maximum flight level sinusoidal vibration inputs are the same for all Delta 2 configurations and are defined at the base of the payload attach fitting. FIGURE : DELTA 2 ACOUSTIC ENVIRONMENT FOR 9.5 ft AND 0 ft0l FAIRING MISSIONS Shock The maximum shock environment at the PAF/Spacecraft interface occurs during spacecraft separation from the Delta 2 (see Figure 2). FIGURE 2 : SHOCK RESPONSE SPECTRUM 372A, 372B, 372C PAF December 2003 Page 3

14 3.4.4 Thermal environment The ascent thermal environments of the Delta 2 fairing surfaces facing the payload, based on historical flight data, are shown in Figures 3 and 4. Temperatures are provided for both the payload fairing (PLF) conical section and the cylindrical section. PLF inboardfacing surface emissivity values are also provided. All temperature histories presented are based on a worstcase trajectory, ignoring expansion cooling effects of ascent. FIGURE 3 : MAXIMUM INTERNAL WALL TEMPERATURE AND INTERNAL SURFACE EMITTANCE (9.5ft FAIRING) December 2003 Page 4

15 FIGURE 4 : MAXIMUM INTERNAL WALL TEMPERATURE AND INTERNAL SURFACE EMITTANCE (0 ft FAIRING, STANDARD AND STRETCHED) December 2003 Page 5

16 3.4.5 Variation of static pressure under fairing As the Delta 2 ascends through the atmosphere, the fairing is vented through a 64.5 cm 2 opening in the interstage and other leak paths in the vehicle. The expected extremes of internal pressure during ascent are presented in Figure 5 for the 2.4 m, 2.9 m and 3 m fairings. FIGURE 5 : DELTA 2 PAYLOAD FAIRING COMPARTMENT ABSOLUTE PRESSURE ENVELOPE Spacecraft compatibility tests Sinusoidal vibration tests QUALIFICATION ACCEPTATION DESIGNATION FREQUENCY RANGE (Hz) LEVELS (0 PEAK) FREQUENCY RANGE (Hz) LEVELS (0 PEAK) Thrust cm double ampl cm double ampl g g Lateral cm double ampl..0 g g Sweep rate 2 oct. /min 4 oct. /min December 2003 Page 6

17 3.5 Operations constraints Ground constraints : The mission Director Center provides the necessary seating, data display and communications to control the launch process. Flight constraints : Launch rate capability : Delta 2 has had an average utilization of approximately one launch every 60 days. Procurement lead time : The spacecraft questionnaire is to be completed by the spacecraft agency at least 2 years prior to launch to provide an initial definition of spacecraft characteristics. 4. LAUNCH INFORMATION 4. Launch site The Boeing Company operates two launch sites within the continental U.S. : Eastern Range (ER) in Florida and Western Range (WR) in California. The Space Launch Complex (SLC) of the ER is located at Cape Canaveral Air Force Station (CCAFS) and consists of two launch pads, designated SLC7A and SLC7B. The SLC2 in the WR is located at Vandenberg Air Force Base (VAFB) and is typically used for missions requiring highinclination orbits, while SLC7 is used for low to medium inclination orbits. Both launch complexes are open to commercial and government customers. Depending on the specific mission requirement and range safety restrictions, the Delta , 7400 and 7900 series vehicle can be launched from either the ER or WR launch site (7900H series can only use the ER launch pad at present). Cape Canaveral Air Force Station a part of the ESMC (Eastern Space and Missile Center), is located approximately 80 km East of Orlando (Florida). This site can accommodate flight azimuths in the range of 65 to 0 deg, with 95 deg being the most commonly flown. Launch Complex 7 consits of two launch pads (7 A and 7 B), a blockhouse, ready room, shops, and other facilities needed to prepare, service and launch the Delta 2 vehicle. Maintenance and launch preparation may be conducted at one pad without impacting operations at the other. LAUNCH PAD LATITUDE (degrees North) LONGITUDE (degrees West) LAUNCHER AZIMUTH 7 A B Launch complex 7 is convenient for orbit inclinations from 28.5 to 5. Commercial spacecraft will normally be processed through the Astrotech facilities. Other facilities at CCAFS, controlled by NASA and USAF, can be used for commercial spacecraft under special circumstances. In addition to the facilities required for the Delta 2 launch vehicle, specialized facilities are provided for the spacecraft: payload processing facilities : NASA provided hangars AO, AM, or AE, Astrotech Space Operations in Titusville, hazardous processing facilities: NASA provided payload spin test facilities: Explosive Safe Aera 60 (ESA 60), Spacecraft Assembly and Encapsulation Facility 2 (SAEF 2), Cargo Handling Storage Facility at KSC or CCAFS, Astrotech Space Operations, Assignment of these facilities is controlled by the respective owners. December 2003 Page 7

18 FIGURE 6 : ORGANIZATIONAL INTERFACES FOR COMMERCIAL USERS The arrangement of SLC7 and an aerial view are shown in the following figures. FIGURE 7 : DELTA CHECKOUT FACILITIES December 2003 Page 8

19 FIGURE 8 : ASTROTECH SITE LOCATION Typical schedules of integrated activities from spacecraft weighing in the HPF (hazardous processing facility) until launch are shown as launch minus (T) workdays. Saturdays, Sundays, and holidays are typically not scheduled workdays and therefore are not Tdays. The Tdays, from spacecraft mate through launch, are coordinated with the customer to optimize onpad testing. All operations are formally conducted and controlled using launch countdown documents. The schedules of spacecraft activities during that time, also called countdown bar charts, are controlled by the Boeing chief launch conductor. Tasks involving the spacecraft or tasks requiring that spacecraft personnel be present are shaded for easy identification. FIGURE 9 : TYPICAL TERMINAL COUNTDOWN (T0 DAY) December 2003 Page 9

20 Vandenberg Air Force Base (VAFB) in California The pad available at VAFB for Delta launches, the Space Launch Complex SLC2, is convenient for orbit inclinations rorm 63 to 45. Flight azimuths in the range of 90 to 225 are currently approved by the 30th Space Wing, with 96 deg being the most commonly flown. LAUNCH PAD LATITUDE (degrees North) LONGITUDE (degrees West) LAUNCHER AZIMUTH SLC The manufacturer maintains an operations team at VAFB that provides launch services to NASA, commercial and USAF customers. FIGURE 20 : VANDENBERG AIR FORCE BASE (VAFB) FACILITIES FIGURE 2 : SPACE LAUNCH COMPLEX 2 VANDENBERG TEST CENTER December 2003 Page 20

21 FIGURE 22 : SLC2 MOBILE SERVICE TOWER/FIXED UMBILICAL TOWER ELEVATIONS The integration process is designed to support the requirements of both the launch vehicle and the payload. The typical integration process encompasses the entire life of the launch vehicle/payload integration activities; Ldate is defined as calendar day, including workdays and scheduled nonworkdays such as holidays. At its core is a streamlined series of documents, reports, and meetings that are flexible and adaptable to the specific requirements of each program. December 2003 Page 2

22 FIGURE 23 : TYPICAL MISSION INTEGRATION PROCESS FIGURE 24 : SPACECRAFT SUPPORT AERA December 2003 Page 22

23 4.2 Sequence of flight events Typical profile for three stage missions is shown in figure 25. FIGURE 25 : DELTA TYPICAL MISSION PROFILE FIGURE 26 : TYPICAL DELTA /7925H MISSION PROFILE GTO MISSION December 2003 Page 23

24 Event VEHICLE CONFIGURATION 7325/ /7925H 7326/ /7926H First stage Main engine ignition T + 0 T + 0 T + 0 T + 0 Solid motor ignition (3,4 or 6) T + 0 T + 0 T + 0 T + 0 Solid motor burnout (3,4 or 6) T + 63 T + 63 or 77 T + 63 T + 63 or 77 Solid motor ignition (3) N/A T + 66 or 79 N/A T + 66 or 79 Solid motor separation (3,4 or 3/3) T + 66 T + 66/67 or 8/82 T + 66 T + 66/67 or 8/82 Solid motor burnout (3) N/A T + 29 or 57 N/A T + 29 or 57 Solid motor separation (3) N/A T + 32 or 60 N/A T + 32 or 60 MECO (M) T T T T Second stage Activate stage separation bolts M + 8 M + 8 M + 8 M + 8 Stage II ignition M M M M Fairing separation M + 39 M + 39 M + 39 M + 39 SECO (S) M + 45 M M M Stage II engine restart S + 60 S S + 60 S SECO (S2) S + 63 S S S + 70 Third stage Activate spin rockets, S S S S Separate stage 2 S S S S Stage 3 ignition S S S S Stage 3 burnout S S S S Spacecraft Spacecraft separation S S S S TABLE 3 : DELTA 2 TYPICAL EASTERN LAUNCH SITE EVENT TIMES (in seconds) Event VEHICLE CONFIGURATION / First stage Main engine ignition T + 0 T + 0 T + 0 T + 0 Solid motor ignition (3,4 or 6) T + 0 T + 0 T + 0 T + 0 Solid motor burnout (3,4 or 6) T + 64 T + 64 T + 64 T + 64 Solid motor ignition (3) N/A T + 66 N/A T + 66 Solid motor separation (3,4 or 3/3) T + 83 T + 86/87 T + 99 or 83 T + 86/87 Solid motor burnout (3) N/A T + 29 N/A T + 29 Solid motor separation (3) N/A T + 32 N/A T + 32 MECO (M) T T T T Second stage Activate stage I/II separation bolts M + 8 M + 8 M + 8 M + 8 Stage II ignition M M M M Fairing separation M + 29 M + 9 M + 29 M + 9 SECO (S) M + 45 M M M Stage II engine restart S + 60 S S + 60 S SECO (S2) S + 63 S S S + 70 Third stage Activate spin rockets, S S S S Separate stage 2 S S S S Stage 3 ignition S S S S Stage 3 burnout S S S S Spacecraft Spacecraft separation S S S S TABLE 4 : DELTA 2 TYPICAL WESTERN LAUNCH SITE EVENT TIMES (in seconds) December 2003 Page 24

25 4.3 Launch record data LAUNCH DATE NUMBER OF SATELLITES ORBIT RESULT REMARK Circular Success GTO Success GTO Success GTO Success LEOcircular Success LEOcircular Success LEOcircular Success GTO Success LEOcircular Success GTO Success LEOcircular Success GTO Success LEOcircular Success LEOcircular Success LEOcircular Success LEOcircular Success LEOcircular Success LEOcircular Success LEOcircular Success LEOcircular Success GTO Success GTO Success LEOcircular Success Interplanetary Success GTO Failure GTO Success LEO Success LEO Success LEOcircular Success GTO Success LEOcircular Success LEOcircular Success Interplanetary Success Interplanetary Success LEOcircular Failure GTO Success LEOcircular Success LEOcircular Success 7925 December 998 Page 25

26 LAUNCH DATE NUMBER OF SATELLITE ORBIT RESULT REMARK GTO Success 7925 ETR GTO Success 7925 ETR LEO Success 7326 ETR GTO Success 7925 ETR Interplanetary Success 7425 ETR Interplanetary Success 7425 ETR Interplanetary Success 7426 ETR LEO Success 7925 ETR LEO Success 7326 WTR LEO Success 7925 ETR LEO Success 7925 ETR LEO Success 7925 ETR LEO Success 7925 ETR Interplanetary Success 7925 ETR GTO Success 7925 ETR Interplanetary Success 7425 ETR Interplanetary Success 7326 ETR Interplanetary Success 7425 ETR MEO Success 7925 ETR MTO Success 7925 ETR Interplanetary Success 7925 ETR Interplanetary Success 7925H ETR MEO Success 7925 ETR LEO Success 7925 ETR LEO Success 7925 ETR Interplanetary Success 7925H ETR LEO Success 7925 ETR Failures: two LAUNCH DATE RESULT CAUSE The satellite had to use its onboard fuel to raise the spacecraft to its proper orbit An anomaly occured approximately 2 s into flight. A selfdestruct was initiated then at 2 s the flight safety officer sent a precautionary destruct signal. The explosive lines used for the separation had been damaged by exposures, causing one of the SRB failing to separate. A 6.3 m vertical rupture occured for unknown reasons of one of the SRB. Reliability Previsional reliability: Success ratio: 96.9% (63/65) 4.4 Planned launches December 2004 Page 26

27 Six launches are planned in 2005 for the 7925 DELTA 2 launch vehicle configuration. In December 2002, NASA ordered 2 Delta II for launches between 2006 and 2009 and 7 more options for launches from Vandenberg. December 2004 Page 27

28 5. DESCRIPTION 5. Launch vehicle FIGURE 27 : DELTA 7925 LAUNCH VEHICLE December 2004 Page 28

29 H0 FIGURE 28 : CONFIGURATIONS OF THE DELTA 2 LAUNCH VEHICLE WITH THIRD STAGE 5.2 Overall vehicle (Delta 27925) Overall length : m Maximum diameter : 2.4 m (without fairing) Liftoff mass : t (7925) December 2004 Page 29

30 5.3 General characteristics of the stages STAGE (a) 3 (b) (Delta 7326) Designation GEM 40 () (GEM H) Star 48 Star 37FM Number of motors 3, 4 or 9 Manufacturer Alliant Techsystems Boeing Boeing Boeing Boeing Length (m) 2.97 (4.7) Diameter (m).06 (.7) Dry mass (t).35 (2.035) (4) Propellant: solid (2) liquid liquid solid (3) solid Type storable nonstorable storable storable storable Fuel HTPBAl RP Aerozine 50 HTPB HTPB Oxidizer NH 4 CIO 4 LO 2 N 2 O 4 NH 4 CIO 4 NH 4 CIO 4 Propellant mass (t): Total Fuel. Oxidizer Water (usable) Tank pressure (bar) Total liftoff mass (t) (max).47 () GEM: Graphite Epoxy Motor, GEM 46 for Delta 27925H (2).3 class (3) TPH3340 type (4) with fixed nozzle, 2.28 t with TVC (ThrustVector Control) (a) Delta The Star 48B motor is used as the third stage of the standard Delta (b) Delta The smaller Star 37FM motor is attached to the satellite and is used to place it in its proper orbit. Instead of nine solid rocket boosters (with six ignited at liftoff), the Medlite configuration used only three groundlit solids (Delta 27325/7326) or four (Delta 27425). December 2004 Page 30

31 Upper part DESIGNATION SPIN TABLE PAYLOAD ATTACH FITTING Manufacturer Mass 839 kg 5.4 Propulsion STAGE (Delta 7326) Designation SSRM RS 27A AJ08 K STAR 48B STAR 37FM Manufacturer Alliant Techsystems Rocketdyne Aerojet Thiokol Thiokol Number of engines Engine mass (kg) Feed syst. type turbopump helium Mixture ratio Chamber pressure (bar) Cooling Specific impulse(s): Sea level Vacuum Thrust (kn): Sea level Vacuum Burning time(s) Nozzle expansion ratio Restart capability 0: no 2: no 65: : no 43.7 : no * SSRM : Strapon Solid Rocket Motor () + Verniers 2x 4.45 kn December 2004 Page 3

32 5.5 Guidance and control 5.5. Guidance Inertial guidance unit on first and second stages + DELCO processor unit Control STAGE 0 (SSRM) 2 3 Pitch, yaw Roll GEM 40 : Stage provides control. SSRMs have fixed 0 deg nozzle cant. GEM 46 : 3 motors have hydraulic gimbaled nozzles with flexible seals, other have fixed nozzle Hydraulic nozzle gimbal 2 Rocketdyne LR0NA vernier engines Hydraulic gimballing Cold gas nitrogen Spinstabilized with hydrazine NCS Precision 6. DATA SOURCE REFERENCE. DELTA II PAYLOAD PLANNERS GUIDE BOEING October DELTA 2 spacecraft users manual 3. AIAA : DELTA 2 a new generation begins July AIAA 89233: GEM for the DELTA 2 July PAMD user's requirements document 6 Aviation Week Space Technology / November 2, 998, page 32 7 DELTA II Payload Planners Guide Update January 2003 The Boeing Company December 2004 Page 32