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Produced by the NASA Center for Aerospace Information (CASI) X-470-69-406 PREPRINT

r - N Asa r^i, x= 63(o 77

Y DELTA-A SCIENTIFIC AND APPLICATION SATELLITE LAUNCH VEHICLE

CHARLES R. GUNN d co N ^(1 o

SEPTEMBER 1969

- GODDARD SPACE FLIGHT CENTER 4D GREENBELT, MARYLAND n ^ ^^ OE 1 ITMRUI

j 'PAGES) CODEI

(NASA CR OR TNIX OR AD NUFIDC RI ICA EGORYI

X-470-69-406 i i

DELTA SCIENTIFIC sN0 APPLICATION SATELLITL LAUNCH Vh::ICLF

Charles P... Gunn

September 1969

GODDARD SPACE FLIGHT CENTER Greenbelt, Maryland f^;ECEp!J^G PAGE

Nor fi rL

ABSTRACT

The Delta launch vehicle, Model 904 is described for potential users. This new model of Delta is composed of the Long Tank Tl.or first stage, thrust augmented by 9 Castor II solid propellant motors; the Delta second stage, converted to higher energy storable propellants, uprated by adoption of the Titan Transtage ablative engine and modernized with a strap-down inertial system that replaces independent autopilots/radio guidance systems in the first and second stages; and the elongated TE-364-4 solid propellant third stage motor that evolved fromthe Surveyor spacecraft retromotor. A brief historical summary of Delta's evolutionary growth and flight record attests to the success of the Delta philosophy of making maximum use of current technology and flight proven components from other space programs. Performance, flight environment, organizational interfaces, spacecraft integration requirements, launch operations and costs are provided.

This new model of Delta is to be available in mid 1971, cost about $5. 0 million dollars and be capable of injecting 4, 000 pounds into low earth orbit, 1300 pounds into synchronous transfer orbit, or escaping 900 pounds of payload.

iii 1" LCECING PAVE ELA,tiK NOT FILMED. CONTENTS

Page

ABSTRACT ...... iii

I. THE EVOLUTION OF DELTA ...... 1

II. THE DELTA LAUNCH VEHICLE ...... 4

A. Vehicle Description ...... 4 B. Flight Sequence and Performance ...... 11 C. Flight Environment ...... 16 D. Organization and Interfaces ...... 23 E. Spacecraft Integration and Launch Operation ...... 25 F. Cost ...... 28

TABLES

Table Page

I Delta Standard Attach Fittings ...... 12

II Delta Performance Capabilities ...... 18

III Synchronous Transfer Orbit Dispersions ...... 19

IV Sun-Synchronous Orbit Dispersions (Two Stage Vehicle) . . . . 21

V Delta Critical Flight Environment ...... 22

VI Delta Launch Costs (1971) ...... 29

v ILLUSTRATIONS

Figure Page

1 Delta Evolutions ...... 2

2 Delta (Model 904) ...... 5

3 Delta Raging Schematic ...... 7

4 TE 364-4 Third Stage and Spintable ...... 10

5 Delta Payload Attach Fittings ...... 13

6 Delta Fairing Spacecraft Envelope ...... 14

7 Delta Flight Sequence of Events for a Synchronous Transfer Mission ...... 15

8 Characteristic Velocity for Delta Model 904 ...... 17

9 Delta Flight Sequence of Events for a Polar Circular Mission ...... 20

10 Organization and Interfaces ...... 24

11 Delta Mission Analysis and Integration ...... 26

12 Delta Payload Cost per Pound Synchronous Transfer Mission — ETR ...... 31

vi DELTA-A SCIENTIFIC AND APPLICATION SATELLITE LAUNCH VEHICLE

I. THE EVOLUTION OF DELTA

The evolution of the Delta launch vehicle, shown in Figure 1, reaches back thirteen years when, in 1955, the United States participated in the International Geophysical Year (IGY) and undertook the development of the Vanguard three- stage launch vehicle; in the same year the Air Force initiated the development of the Thor IRBM. With modifications, the Thor became the first stage of Delta; the Vanguard second stage propulsion system, evolved through the Able programs, became the Delta second stage propulsion system; and the Vanguard X-248 third stage solid propellant rocket motor was adapted as the third stage for Delta. The development and integration of these systems and the production of twelve (12) vehicles was started in early 1959 under prime contract to the Douglas Aircraft Company. The initial objective of the Delta program was to provide an interim space launch vehicle capability for the medium-class payloads until more sophis- ticated vehicles as Scout and Agena, then under development, could be brought to operational status. The development program spanned 18 months. In a little over two years, following the development period, eleven of the twelve vehicles were launched successfully carrying, among others, the first passive communications satellite, Echo I (August 1960), the cooperative NASA/United Kingdom Ariel I (April 1960), the TIROS II through VI series, the first Orbiting Solar Ob- servatory, and the first private industry satellite, AT&T's Telstar I (July 1962). The total development cost, including the twelve vehicles (Model DAI-19) and launch support, was approximately $43, 000, 000, compared to the $40, 000, 000 estimated at the outset of the program.

Before the development program was complete the number of missions planned for Delta outstripped the interim buy of twelve vehicles, so an order was placed for fourteen additional vehicles. This follow-on buy of Deltas (Models A and B) incorporated lengthened second stage propellant tanks, a higher energy second stage oxidizer, transistorized guidance electronics, and assiduous application of high-reliability semiconductors in flight critical circuits. This model of Delta carried NASA's first active communications satellite, Relay I (Decem- ber 1962), and the first synchronous satellites, Syncom I and II (February and July 1963).

The next production order of Deltas (Models C and D) in 1963 brought the adaption of the USAF developed improved Thor booster with thrust augmentation provided by three strap-or solid propellant motors and the adaption of the Scout developed X-258 to replace the X-248 third stage motor. The first thrust augmented Delta (TAD) carried Syncom III (August 1964), the first equatorial

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2 synchronous conunumications :satellite. The second TAD vehicle orbited tho first commercial communications satellite, Comsat Corporation's Early Bird Satel- lite (April 1965).

Another order of Deltas in 1964 brought the development of the Improved Delta (Model F) . The Improved Delta model adapted and extended the large diameter propellant tanks from the Able-Star stage, and thereby nearly doubled the propellant capacity of the previous Delta second stage. The larger diameter tanks in addition permitted adaption of the five foot diameter Nimbus fairing developed for the USAF Agena stage. Improved Delta also adapted the USAF developed FW-4 solid propellant motor to replace the X-258 third stage motor (Model E1). The first Improved Delta was launched November 1965 and among the missions carried on this model of Delta are the near polar Geophysical Or- biting Satellites, GEOS A and B; the lieliocentric Pioneer series A through D; the low earth orbiting Biological Satellite, BIOS A through C; the synchronous communications satellites, Intelsat F1 through F4; the lunar orbiting Anchored Interplanetary Monitoring Probe, A-IMP A and B; the sun-synchronous ESSA 2 through 6; the High Eccentric Orbiting Satellite, HEOS developed by the European Space Research Organization and the C uiadian International Satellite for Iono- spheric Studies, ISIS.

In 1966 Delta undertook to adapt the Surveyor spacecraft solid propellant retromotor as a new third stage. The spherical case was modified to mate to a spin-table assembly and the motor, designated TE-364-3, was requ.dified for the Delta spinning environment. The first Delta using this third stage motor, Delta Model J, was launched in July 1968 and carried the Radio Explorer, RAE-A spacec raft.

At about the sane time Delta initiated the adaption of the TE-364-3 motor, the USAF undertook the uprating of the Thor booster by lengthening the liquid oxygen and RP-1 fuel tanks and converting the fuel tank to a constant 8 foot diameter. This Long Tank Thor carries about 47 percent more propellants than previous models. In September 1968. D^lta launched its first Long Tank Thor %p ith the Improved Delta second stage and TE-364-3 third stage. The Delta Model M carries, among others, the Intelsat III series and the British Skynet and NATO communications satellites.

In early 1968, Delta started a redesign of the Long Tank Thor engine section to permit the addition of a second set of three thrust augmentation solid motors. The first Delta Model 116 with six solid motors is to be launched from the Western Test Range late this year and carry the NASA TEROS Operational Satellite, TOS- M into a 800n. mi. circular swi-synchronous orbit. The Delta Model M6 is to be used also to carry the Interplanetary Monitoring Probe I, the RAE-B, and European Space Research Organization TD-1 spacecraft.

3 To date, Delta is launching over fifty percent of NASA's unmanned space- crafts each year and has been selected for the use of private industry and foreign governments. The reliability and cost effective history of Delta is, in a large part, attributable to the technical approach taken at the outset of the program and still adhered to tcxlay. This approach is to use current technology raid lliglit proven components wherever possible from other space programs. The result- ant vehicle is normally heavy, but cheap and has a high probability of performing repeatedly and reliably from the outset. Delta has never considered it necessary to have a pre-operational or development fli ght test lawuh for any of the ten major changes made to the vehicle. And with the exception of the first Delta launch in 1960, there has never been a failure of the first flight article on its maiden launch. The criteria for evaluating improvements to Delta is that, they must meet the mission requirements at the lowest possible cost and risk without compromise of Delta's reliability record—currently 66 successes out of 73 launches. For this reason Delta has wherever possible, adapted flight proven components from other programs. The next evolutionary uprating of Delta is consistent with this past pattern of change.

H. TIIF DELTA LAUNCH VEHICLE

To keep pace with the growing launch capability requirements of scientific and applications satellites, both domestic and foreign, the Delta launch vehicle is being uprated in performance capability, in guidance accuracy, and modernized to enchance overall systems reliability. This new Delta with nine strap-on solid motors (Model 904) is shown in Figure 2 and is composed of the Long Tank Thor, now with an option of thrust augrr►entation from combinations of 3 to 9 strap-on solid motors; the Delta second stage converted to higher energy storable propellants with the substitution of the Aerojet Titan Transtage ablative engine for the current Vanguard/Able-Star vintage regenerative engine, an on-board strap-down inertial measurement unit and general purpose navigation computer in the second stage to replace the first and second stage autopilots and the Western Electric Company (WECO) radio-guidance system; and the Thiokol TF.- 364-4 solid propellant third stage motor. This Delta is described together with its performance, flight environment, spacecraft integration organization, mile- stones and operations and launch costs. The Delta Model 904 is now scheduled to launch the Planetary Explorer series, the Interplanetary Monitoring Probes J and K, and the Synchronous Meterological Satellites.

A. Vehicle Description

The three st: ge Delta vl,hicle Model 904 shown in Figure 2, stands 106 feet and weighs 261, 000 pounds at lift-off. The vehicle is designed for ascent through

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The first stage liquid propellant core is the Long Tank Thor. The core is M S feet in dameter, 60 feet long, carries 147, 000 pounds of RP-1 and liquid oxygen propellants and is powered by a turbopump fed Rocketdyne main engine that develops 17.5, 000 pounds thrust at lift-off. The core burns to propellant depletion about 220 seconds after lift-off (T+220) at an altitude of 60 to 70 nautical Guiles (n.mi.). Thrust augmentation solid propellant riotors attach at the base of the first stage core to an universal boat-tail (UBT) engine section structure. The UBT is structurally designed and thermally insulated to carry up to 9 Thiokol Castor II solid motors (TY-354-5) or 3 Algol class of solid motors with 6 Castor II's. Use of the Algol class of solid motors is not approved at this time. The UBT uses the same solid motor attach and separation hardware as on all Thrust Augmented Thors . The new Ignition and Separation Sequence Unit developed to ignite and separate the solids after lift-off draws directly from the present unit. The simple, straight forward structural, plumbing and electrical modifications to the present engine section obviates the need for extensive testing or a pre- operational flight test.

Normally, the thrust augmentation motors are built-up in sets of 3. Up to 6 motors can be ignited on the pad and the remainder no sooner than 38 seconds after lift-off in order to hold the vehicle acceleration induce l loads on the propellant tank bottoms within allowable limits. The Castors II motors each develop 33, 000 pounds thrust at ignition, burn for 40 seconds and are jettisoned from the core in sets of three at five seconds intervals starting at T+85. This time is dictated by considerations of combined dynamic pressure angle-of-attack loadings on the jettison mechanism aM a Range Safety requirement for an off- shore impact of the expended motor:.. Jetti;•on is effected by firing and explosive bolt holding a clamped ball-socket joint. Acceleration of the core plus aerodynamic drag on the motors eject the cases away from the veh i cle as shown in Figure 3.

During powered flight pitch and yaw steering is exerted by gimballing the core main engine. Roll control is effected by differentially gimballing a pair of small outboard vernier engines. Subsequent to main engine shut-down the ver- niers coi.tinue tc operate for about 12 seconds, damping shut-down transient and stabilizing the vehicle for staging of the second stage.

Guidance and control of the first stage originates from a new strap-down inertial guidance --Ystem (SIDS) located in the second stage. An inertial measurement unit composed of an orthognal set of linear accelerometE•:s and rat

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7 ( '^ integrating gyros sense 1 .chicle linear and angular motion. An on-hoard general purpose storable program computer directs the vehicle through a pre-programmed trajectory, calculates the present position and velocity, then integrates the nominal trajectory forward to determine the position and velocity at injection. This state vector is then compared to the desired termir -d condition to determine the required steering commands to reach the desired terminal injection point. This computer also performs the functions of commanding discreet flight functions as well as issuing pre-programmed command attitude rates during powered and coast guidance phaze of flight. The new SIGS replaces two independent first and second stage autopilots on Delta and the associated Western Electric Company (WECO) ground radio-guidance system. With SIDS, the relatively expensive ground tracking station support costs are avoided and flight proP'_e and attitude constraints together with horizon viewing limitations are eliminated. Thus, launch costs are reduced, the vehicle performance is increased and closed loop guidance is extended beyond the radio horizon with an attendant increase in injection accuracy. The introduction of SIDS also afford a modernization of the total Delta vehicle guidance and control system which today still utilizes acuum tubes, punched tape mechanical programmers, mechanical rotary inverters, and solder type electrical connectors.

The interstage section between the first and second stages is provided with exhaust and access ports to allow the escape of exhaust gases during second stage engine start in the interstage. These ports are covered in flight with band assemblies that are jettisoned (Figure 3) at first stage main engine shut-down. Four seconds later explosive bolts that attach the two stages are fired and the second stage engine is started.

The Delta second stage is 17 feet long, approximately 5 feet in diameter and weighs 12,600 pounds at ignition. The Aerojet engine developed for Vanguard and adapted by Delta is now replaced by the flight proven Aerojet engine developed for the Titan III Transtage. This pressure fed ablative and radiation cooled engine develops 9460 pounds thrust at an uprated 125 pounds per square inch chamber pressure and operates for about 320 seconds on Aerozine-50 and N,04 storable propellants. To adapt the engine to Delta the gimbal mount is redesigned to accept the Delta actuators and the radiation cooled nozzle extention is trimmed back to an area ratio of 26 to 1 from 40 to 1 to keep the same dimensional envelop for staging clear-inces .

The common bulkhead between the fuel and oxidizer tanks i - shifted slightly aft io adjust .o the new propellant mixture ratio and the helium pressurization revelator for the tap ks is mcdified redu,_:e Vie press • rP, tv iaawn the engine requirements. Si nce the engine, designated AJ 10-118F for Delta use Is led through a single bi-prore_'.ant valve the repeatability and re,'i,:;)il Ity of the start

-i and stop sequence is considerably higher than the current Delta engine with independent pilot and main fed valves. Hence, the overall reliability of the second stage propulsion system is enhanced and at the same time the stage performance is boosted about 3 percent.

During the second powered flight, pitch and yaw steering is provided by gimballing the engine and roll is controlled by cold nitrogen gas jets. Cold nitrogen gas jets control the vehicle in all axes during coast and provides propellant settling ullage thrust for restarting the engine. The control system electrical power and nitrogen gas supply is capable of maintaining second stage attitude for a little over two hours. For long second stage coast periods before third stage spin-up and separation the second stage may be reoriented with respect to the sun or the vehicle placed in a slow yawing or pitching tumble to al- leviate assymetric solar heating of the spacecraft.

Peripheral second stage systems include a "C" band tracking beacon, a PDM/FD1/FM 4:5 x 20 telemetry system. dual command destruct receivers and associated power supplies. The command destruct receiver system is capable of both cutting-off the second stage engine or destroying an errant vehicle.

The third stage assembly consists of a spin-table, the Thiokol TE-364-4 solid propellant motor, spacecraft attach fitting, spacecraft and the spacecraft fairing. The spin-table shown in Figure 4 consists of a bearing support structure and a conical third stage motor pedestal truss that is divided into four petals hinged at the base and clamped to the equator of the third stage motor by a "V" band. The "V" band is held in tension by two explosive bolts that are fired two seconds after the motor and spacecraft are spun-up and the 15 second time delay squib that ignites the TE-364-4 motor is started. The released petals fly out- ward under centrifugal force, releasing the third stage from the spin table (Figure 3) . At the same instant the second stage is backed away from the free spinning third stage by venting propellant residual pressurant (helium) over- board through two retrojets. Approximately thirteen seconds later the third stage motor is ignited by the time delay squib. The TE-364-4 motor is essentially identical to the -3 model previously used except that a 14 inch cylindrical section is added between the two hemispherical halves of the case. The propellant weight is increased to 2350 pounds from 1440 pounds, it burns for 44 seconds and develops an average thrust of 15, 000 pounds.

Torque to the spin table is imparted by combinations of small solid propellant rocket motors, which provide spin rates from 30 to 100 rpm (f10 percent) for spacecraft roll moment of inertia ran ging from 20 to 170 slug-feet squared. A lower limit of approximately 30 rpm is dictated by minimum dynamic stability of the third stage/spacecraft assembly during third stage motor burning. If less

9 ATTACH F11 MOUNTS

SEPARAT 10 CLAMP

QUARTER PANELS (4) S SUPPORT STR U CTU R ZING

SE __.._ _...^_ STRUCTURE Figure 4. TE 364-4 Third Stage and Spintable than 30 rpm is desired the effect upon orbit injection errors must be carefully a.ssossed. The anticipated maximum spin rate users would desire was 100 rpm, consequently the third stage motor is qualified only up to this spin rate.

The spacecraft is clamped to the attach fitting by a circular V-clamp-band assembly that releases by firing two explosive bolt cutters subsequent to third stage motor burn-out. Separation from the expended third stage is then effected by a separation spring, or springs, which provides the spacecraft with a relative separation velocity of 6 to 8fps with respect to the expended third stage motor. Although peculiar spacecraft requirements may dictate the design of a special spacecraft attach fitting, a number of standard Delta fittings are avilable. These

10 IE

are specified in Table I and shown in Figure 5. These fittings use either a small rocket or yo weight system to tumble the expended third stage motor after spacecraft separation to preclude possible motor outgassing from accelerating it into the spacecraft. Also available is a yo-yo weight despin system which can despin the third stage and spacecraft combination prior to spacecraft sepa~ation. At- tach fittings include timer assemblies, battery and delay squib switches. The timers are initiated by the second stage computer and run on mechanical energy until reaching a predetermined time to fire the spacecraft separation clamp- band bolt cutters and a pair of squib switches. Two seconds later the squib switches initiate a small rocket or yo weight to tumble the expended third stage motor.

For users requiring real time third stage motor performance, environmental or velocity increment information an "S" band telemetry system and a "C" band tracking beacon are developed and flight proven. These are carried on either the spacecraft attach fitting or on the third stage motor as optional equipment.

The spacecraft fairing is fiberglass, and constructed in two-half-shells that are brought up around the spacecraft laterally and clamped together by three strap assemblies that are released in flight by explosive bolts. Spring cartridges thrust the half-shells laterally and pivots at the base of the fairing cause the shells to rotate rearwards and clear the vehicle (Figure 3) . Normally, the fairing is jettisoned within 5 to 20 seconds after second stage start. Fairing jettison time is dictated by the free molecular heating rate that can be tolerated by the spacecraft. Normally, the heating rate is held below 0.1 BTU/Ft 2 -sec. or about equivalent to the solar heating rate to the spacecraft. Aerodynamic heating of the fairing is controlled by application of ablative materials to hold the fairing internal temperature to below 450'F. This precludes any possibility of spacecraft contamination from outgassing of the fiberglass phenolic.

Access ports through the fairing are provided at the locations that meet the needs of the vehicle user. The available fairing internal envelop is shown in Figure G.

B. Flight Sequence and Performance

The Delta flight profile and sequence of events for a three stage synchronous transfer mission having a perigee altitude of 100 nautical miles (n. mi.) an apogee altitude of 19, 400 n. mi. and an inclination of 28.5 degrees is shown in Figure 7. The vehicle is launched from ETR on an azimuth of 95 degrees. The pitch program consists of five discrete first stage pitch rates, two second stage rates, and a coast phase rate. The first and second stage have sufficient perform.uice

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20 x 30 CONICAL ATTACH FITTING 31 x 31 INCH CYLINDRICAL ATTACH FITTING WT: 24-30 LB WT: 54 LB Figure 5. Delta Payload Attach Fittings to place the third stage assembly into a 100n. mi. parking orbit. After second stage engine cut-off, the stage then coasts to a point just short of the Equator where third stage spin-up, separation and ignition occur. The third stage burns out directly over the Equator at an altitude of 100 n. mi. , an inertial flight path angle of zero degrees, and with sufficient velocity to coast the spacecraft to an altitude of 19,400 n.mi. on the opposite side of the Earth so that the line of apsides lies in the equatorial plane to permit the spacecraft apogee motor to rotate the transfer orbital plane into the equatorial plane as part of the circu- larization manuever.

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Payload weight versus characteristic inertial velocity for Delta from the Eastern Test R:ulge in Florida and the Western Test Runge in California is shown in Figure 8. The performance capability for a number of scientific cool applications missions carried on Delta is summarized, in Table I1. These Delta performance capabilities are the useful load that can be carried above the last powered stage and thus includes the spacecraft weight and its attach fitting hardware weight. The first number in the Delta model designation (3. G, or 9) in- dicates the number of thrust augmentation solid motors, the second digit (0) the Delta second stage with SIGS and the AJ 10-118F engine, the third digit (4) notes the -4 version of the TE-364 third stage motor.

The injection accuracy of Delta is strongly dependent on the trajectory profile. The single largest injection error source is the unguided, spin-stabilized third stage. Nearly two thirds of the errors in injection velocity :und attitude are caused by dispersions in the motor tote impulse and lateral tip-off impulses applied during separation from the second stage and at motor ignition. Typically the 99 percent probability dispersions for nominal 100n. mi. by 19,400n. mi. synchronous transfer trajectory is shown in Table III.

Though these dispersions are moderately large, it should be remembered that a synchronous communications satellite must carry a propulsion system for station keeping and hence the penalty paid in additional propellant to trim out injection errors is quite small.

The flight profile and sequences of events for a two stage (Delta; Model 90) sun synchronous 500n. mi. circular orbit mission is shown in Figure 9. The vehicle is launched from WTR on an azimuth of 193 degrees. The second stage is injected into a 100 by 500 11 - mi. Hohmann transfer and coasts 180 degrees to apogee where the second stage restarts and circularizes the orbit. The orbit dispersions for this type of mission is shown in Table IV for both the SIGS and the old autopilot/radio-guidance system that SIDS replaces.

C . Flight Environment

The spacecraft environment is estimated from previous flight measurements. A summary of the expected environment for both the two and three stage Delta vehicle is provided in Table V.

At lift-off the spacecraft is subjected to both lateral wid longitudinal sinu- swlial vibration that local the spacecraft structure dynamically. At the time the three stage Delta lifts off the. launch pins mid the umbilicals are simultaneously retract at statiors ,long the length of the vehicle, the spacecraft can experience a maximum of f2.5g, zero-to-peak (0-1 1), in the vehicle lateral modal fre- quencies (2 to 1311z). Superimposed at this time is a f1.5g (O-P) longitudinal

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E Table II Delta Performance Capabilities

DELTA PERFORMANCE CAPABILITY POUNDS MISSIONS FLIGHT MODE MODEL NUMBER j 304 604 904

L-BI -OSATELLITE TWO STAGE. RESTART 28CO 3400 3700 200 N.M CIRCULAR 100 } 200 N M INCL. = 28 DEGREES HOHMANN TRANSFER.

EARTH RESOURCES TWO START, RESTART 1550 1950 2200 500 N M CIRCULAR 100 • 50C N M. SUN SYNCHRCNO(IS HOHMANN TRANSFER INCL. = 99 DEGREES

IMPROVED TIROS TWO STAGE, RESTART 1250 1550 1800 OPERATIONAL SATELLITE 100 - 800 N M (ITOS) 80G N M HOHMANN TRANSFER. CIRCULAR SUN SYNCHRONOUS INCL = 102 DEGREES

NATO A 'HREE STAGE 1000 1200 1300 SYNCH TRANSFER DIRECT ASCENT. 100 . 19,400 N M SECOND STAGE PLACED INCL. = 28.5 DEGREES IN 100 NM PARKING ORBIT, I PLANETARY EXPLORER THREE STALE 625 725 800 VENUS TYPE II DIRECT ASCENT. C3 = 8.231 KM2/SEC 2 SECOND STAGE PLACED IN 100 N M PARKING ORBIT

13eps oscillation. These combined lift-off oscillatirns typically last for two to five seconds with the peak acceleration lasting one to two cycles. During the la '^t twenty seconds of first stage flight, the Thor exhibits a 20 Hz "pogo" longitudinal oscillation that builds up to f4 g (O-P) at the time the steady stage longitudinal acceleration has reached about 6.3 g. The maximum first stage steady state acceleration of 8 g's is the highest imposed by the two stage Delta. For three stage Delta, the maximum steady state acceleration is lictated by the TE- 364-4 third stage and reaches 23 g's for a 500 pou_.d spacecraft or A g's for a 1500 pound spacecraft.

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22 Random vibration measured at the third stage attach fitting and spacecraft, show power spectrum densities between 0.001 and 0.02g ' /Hz from 2011z to 20001Iz in both lateral and longitudinal axes. The principle source of random acceleration is bowldary layer turbulance over the fairing and the reverse slope of the second stage guidance compartment that excites the structure and feeds up through the third stage assembly to the base of the spacecraft. Acoustical excitation also contributes to the random levels experienced.

At lift-off and transonic, the overall acoustical level inside the fairing is approximately 140db (referenced to 0.0002 dynes/cm'`) from 37.5 to 9600 cps. These levels are present for about 10 seconds at lift-off and again for about 15 seconds at transonic.

Shocks occur at main engine start, thrust augmentation solid motors ignition and jettison, staging, fairing jettison, and spacecraft separation from the expended third stage. For three stage Delta, cutting the bolts to separate the spacecraft from the expended third stage imposes the most severe shock spectrum on the spacecraft. The third stage motor and spin table assembly act to absorb the 'sigh frequency excitation from other sources. Cutting the separation bolts results in an estimated shock spectrum equivalent to a one-third milli- second, 1400 g terminal peal: sawtooth input.

D. Organization and Interfaces

Delta users iiiteriace organizationally with three elements within NASA. This is best illustrated by the relationship that existed between the European Space Research Organization's, Project HEOS, and the NASA Delta Project a,j,l is shown in Figure 10. Agreement between NASA and a foreign space organization for a launch and associated services is established at NASA Headquarters level. This is normally done with a Memorandum of Understanding outlining the principles under which such arrangements are to be made, followed by a specific contract for each mission. The agreed-to policies and fiscal arrangements are then passed through the NASA Office of Space Sciences and Applications (OSSA) Delta Programs Office to the Goddard Space Flight Center (GSFC) Delta Project Office for implementation. The Delta Project Office is vested with the authority and responsibility for carrying out all aspects of a Delta vehicle mission. The Delta Project works directly with the Spacecraft Project to develop and define the spacecraft/vehicle mission requirements, integrate the spacecraft to the vehicle, establish schedules, and determine the final flight readiness of the vehicle. The Delta Project contracts with and directs a single industrial contractor, McDonnell Douglas Astronautics Corporation (MDAC) for the vehicle hardware, mission analysis, and launch support services. Direction of the launch support services furnished by MDAC at the launch site is delegated to the

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24 NASA Kennedy Space Center (KSC). The KSC Delta Operations Team works directly with the Spacecraft Project at the launch site to insure required Range and contractor services are provided and to coordinate the launch site vehicle and spacecraft activities.

This simple organizational structure with short and direct authority and communications lines is a significant factor in the flexibility and responsiveness Delta provides its users.

E. Spacecraft Integration and Launch Operation

Delta vehicle interface constraints together with performance and accuracy estimates are provided to potential vehicle users as soon as the concept of the mission is outlined to the NASA, Goddard Space Flight Center (GSFC) Delta Project Office. The Delta Project welcomes and encourages early definition of prospective missions by potential users. In some instances, mission definition and integration planning has preceded actual mission commitment by two and three years. Experience has demonstrated that this advance and continuous co- ordination between the user and the Delta Project during the period of developing mission requirements, enhances the visibility of both parties and reveals prob- lem areas before final definition of the spacecraft/vehicle interface and trajectory parameters. In general, spacecraft/vehicle planning for new missions follow the pattern and time frame outlined in Figure 11 and starts about one year (T-52 weeks) before launch when the Spacecraft Project provides the Preliminary Mission Definition and Requirements to the Delta Project Office. This definition encompasses the preliminary spacecraft configuration, mass properties, trajectory, and orbital requirements necessary for preliminary vehicle performance evaluation and analysis. A preliminary trajectory with attendant injection error studies and thermal studies is completed within ten weeks. With this visibility, the Delta Project and the Spacecraft Project jointly develop a Final Mission Re- quirement specification (T-40 to T-26 weeks) that includes such constraints as spacecraft orbital lifetime, apogee and perigee altitude and geocentric location, permissible injection errors, injection attitude orientation, launch window criteria, tracking and data retrieval requirements, spacecraft mass properties, and all other data necessary for the preparation of the Final Mission Analysis.

The Spacecraft Project reviews the final mission trajectory about T-35 weeks and the final injection and orbital error analysis about T-23 weeks . The trajectory includes all technical data defining the flight mode, sequence of flight events, vehicle weights and propulsion system characteristics, tabulations of trajectory parameters, weight history, radar look angles, and instantaneous impact loci. Final definition of the maximum and minimum allowable spun rate, apacecraft RF systems, and permissible inflight thermal inputs are provided to Delta by T-26 weeks. A full scale compatibility drawing based on the Spacecraft

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26 Project's final configuration drawings is prepared normally at T-16 weeks. This drawing is primarily to show all clearances between the spacecraft and fairing, attach fitting, and third stage motor and locate the orientation of such features as umbilical connectors, access ports through the fairing, and any special interface wiring between the attach fitting and spacecraft. A Spacecraft Handling Plan is jointly developed and finalized about T-S weeks and describes all hazardous systems, spacecraft test procedures, and details pre-launch work schedules. Typically the spacecraft arrives to the launch site two weeks before launch (T-2) and is built up on the third stage motor assembly the following week and mated to the vehicle on the p

The spacecraft must be statically and dynamically balanced prior to receipt at the launch site. The allowable spacecraft center-of-gravity offset and principle a.^is misalignment is 0.015 inches and 0.002 radians, respectively. For missions where injection attitude is extremely critical for mission success, a third stage assembly composite spin balance is conducted at the launch site.

Delta conducts latuiches from both ETR and WTR. Prograde missions with orbital inclinations of 30 degrees or less are normally launched from ETR and near-polar or retro-grade missions from WTR, though near polar missions have been launched from ETR.

Facilities for the spacecraft Project use at the launch site include spacecraft assembly and checkout laboratories, telemetry, fabrication and cryogenic laboratories, clean rooms, shops, storage, and offices.

The first and second stage mission modific.-tions to the vehicle are made in the MDAC production area in Santa Monica. BeiLore shipment to the launch site, the stages undergo individual and then composite all-systems checkout, exclu- sive of the third stage assembly. The first and second stages are delivered directly to the launch pad, erected, and again undergo systems testing. The thrust augmentation solid motors and third stage solid motors are stored and prepared at the launch site. The thrust augmentation solid motors are mated to the first stage on the launch pad about two weeks before launch. The third stage motor is built up on the spin table and the spacecraft mated with the assembly at about the same time. The spacecraft/third stage assembly is trans- ported in an environmentally controlled canister to an environmental room on top of the mobile services tower around the vehicle and there the assembly is mated to the vehicle. While the spacecraft is mated to the vehicle, spacecraft and vehicle checkout and testing is interspersed and whenever possible to ac- j commodate the spacecraft requirements.

27 On pad checkout of the vehicle culminates in a pre--countdown sin elated flight without propellant on-board, wherein all systems of the vehicle are exercised as they are during the mission. The simulated flight test takes place one week before launch and is followed by final preparation of the vehicle for launch and then a three day countdown to lift-off. If necessary, complete access to the spacecraft can be provided up to four hours prior to lift-off, though normally the fairing is installed about 12 to 16 hours prior to launch. Provisions to contin- uously power and monitor the spacecraft from the blockhouse is provided through the vehicle wiring. While the spacecraft is on the vehicle, thermally and her- metically conditioned, filtered air is provided to the spacecraft right up to 1 ift-off. .

Launches off the same pad at ETR have been conducted at two week intervals, though four weeks is a normal one-shift per day operation. For follow-on spacecraft in a mission series a called-up launch can he made on 90 days notice at no increase in launch costs provided it is an identical mission, the mission peculiar hardware (attach fitting, etc.) have been provisioned and the launch does not impact another scheduled mission. Call-up time may be reduced to 60 days at a cost of about $100, 000 for factory and launch checkout overtime or to ,10 days, provided the vehicle has been previously configured for the mission and completed factory checkout in anticipation of call-up. The 30 day option, however, requires commitment of about $175, 000 of non-recoverable funds if call-up is not exercised. Delta has already launched two cal-up missions in 1969. Al- though the normal cal-up notice requirement was 90 days, these missions were accommodated in 60 and 21 days, respectively. Based on 9 years of experience at ETR, the probability of launching in a window 15 seconds w-:de on a given day is 69 percent. The probability for a thirty minute launch window, typical for most missions, is 86 percent. Recently, when spacecraft development delays impacted a number of Delta launches of lower priority and a call-up launch option was suddenly exercised, the Delta Project launched three spacecraft at one-week intervals to relieve a congestion of six spacecraft awaiting launch. Normally, however, the scheduled launch date desired by the Spacecraft Project can by met.

F. Cost

The projected cost of Delta Model 904 reimbursable launches in 1971 from ETR is about $5.0 million dollars. This includes hardware, trajectory soft- ware, spacecraft integration, launch support services, and :NASA administrative charges. This cost does not include, however, charges made by the U . S. Air Force for Range use that world include trackinb, data acquisition, technical operation: and U. S. Air Force support charges as these costs are highly dependent on mission requirements. A breakdown of costs are provided in Table VI

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29 i ,m,1 are based on actual or estimated expenses billed to outside agency users such as ESSA, Comsat, and ESRO for reimbursement to NASA and a projection of these costs into the 1971 time fra-nc when the Delta, Model 904 shall be available for launch. Actual charges for any given mission will, of course, vary to reflect thc: specific mission requirements.

For launches conducted for outside government agencies and private industry, identifiable launch service charges are segregate(, arc' charged directly against the mission. Indirect or cost not identifiable to a peculiar mission are prorated normally over the duration of a launch services contract or a nu—Iler of Delta launches and allocated accordingly.

Since the first Delta launch in 1960, launen costs have increased from about $2.5 million dollars to $5.0 million dollars, while the performance capability into synchronous transfer orbit, for example, has incr( ased from about 100 pounds to 1300 pounds. The net cost per pound in or►-it then has decreased from $25, 900 to $3, 300 as shown in Figure 12. This cost effective pattern has been accomplished through ten major uprating in the Delta vehicle without a single development flight and in a period when materials and labor cost have more than doubled. Of primary importance to Vie user however, is this growth has peer. accompanied with a demonstrated Plight reliability record of 90 percent. In c-stablishing this figure, Delta also established a record of 22 consecut-1ve suc- cessful space vehicle laimches and again broke this record with another string of 25 successes.

In summary, the Delta, Model 904 that is to be available in mid 1971 will afford users greater performance and mission cost effectiveness together with greater orhit injection accuracy. This new model of Delta shall continue to meet the bulk of domestic and foreign government and industry necks and maintain the availability of a reliable, economical applications and scientific launch vehicle for future missions.

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