Pioneers in Propulsion—A History of CSD, Pratt & Whitney’s Solid Rocket Company

by Charles A. Chase (1)

Abstract

A small group of scientists and engineers, with the support of United Aircraft, created a company which would eventually become one of the world’s leaders in solid propulsion systems. This company would be known for its engineering excellence, pioneering design ideas, systems integration expertise and outstanding levels of dependability for its propulsion systems. This San Jose, California propulsion company, Chemical Systems Division (CSD), became a critical contributor to the space and defense programs of the United States. Early work involved building the Titan-IIIC boosters and eventually included other boosters and upper stages for space applications as well as boosters and upper stages for missiles. Because of its important systems integration capabilities, CSD, through its subsidiary United Space Boosters Incorporated (USBI), also became involved with the integration of the Space Shuttle Boosters. CSD was instrumental in the evolution of propellant and SRM component technologies which significantly increased both propulsion system performance and dependability. CSD also did extensive research and development of hybrid and ramjet propulsion systems.

Founding of a New Propulsion Company In the late 1950s, two forward thinking scientists/engineers developed a plan for improving the propulsion capabilities of the United States. They were Mr. Barnet Adelman and Dr. David Altman. On 1 October 1958 (the same day NASA was established), a small group, led by Adelman/Altman formed the United Research Corporation of Menlo Park (with its first office in Los Angeles, then Palo Alto, CA) to advance and develop liquid and solid rocket propulsion systems, with primary emphasis on solids. This organization was funded by and became a part of United Aircraft (2) (UA) which has since expanded into the large conglomerate known as United Technologies Corporation (UTC). Two other key leaders of this fledgling company were Lt. Gen Donald Putt and Mr. Herbert Lawrence. Gen. Putt was retired from the Air Force where his most recent position was Director of the Air Research and Development Command (ARDC). He joined the group as their president, providing important leadership skills during the critical early days of the company. Herb Lawrence was a very important addition to the engineering prowess of the new company.

Mr. Barnet Adelman Dr. David Altman Lt. Gen. Donald Putt Mr. Herb Lawrence (1) AIAA Fellow, Employee of CSD for 42 years (2) United Aircraft consisted of Pratt & Whitney, Sikorsky, Norden and Hamilton Standard

1

46th AIAA/ASME/SAE/ASEE Joint Propulsion Conference & Exhibit25 - 28 July 2010, Nashville, TN

AIAA 2010-6909

Copyright © 2010 by the American Institute of Aeronautics and Astronautics, Inc. All rights reserved.

What’s in a Name? The United Research Corporation of Menlo Park name was soon changed to United Technology Corp. (UTC), a Subsidiary of UA. After UTC won the Titan IIIC booster program, the name was changed to United Technology Center (UTC), a Division of UA.

Futuristic logo of UTC, the propulsion company Eventually UA made diverse acquisitions, such as Otis Elevator and Carrier Air Conditioning, and thus needed a more diverse sounding name. In 1975 UA borrowed the name of the San Jose Group and changed United Aircraft Corporation to United Technologies Corporation (UTC). The San Jose group was given the new name of Chemical Systems Division (CSD). Ultimately, San Jose experienced one more name change. Pratt & Whitney (P&W), who is world renowned for their jet engines, formally established a space division called P & W Space Propulsion Systems with two groups: West Palm Beach for liquid rockets and San Jose for solid rockets. For simplicity, this paper will use the acronym CSD as the name of the San Jose company.

Pioneering Large Solid Rockets CSD was primarily interested in solid rockets—large solid rockets. Prior to the establishment of CSD, solid rockets were limited in size because of two primary factors: 1) propellant physical properties and 2) shipping limitations because they were of monolithic construction. Key elements which allowed CSD to move forward and get beyond these limitations were: 1) a patent by Adelman to segment very large solid rockets, 2) a patent by Lawrence concerning how segments could reliably be connected to each other and 3) the development, under the direction of Altman, of a Polybutadiene/Acrylic Acid/Acrylonitrile (PBAN) propellant system (which provided the needed physical properties). In February of 1960, CSD broke ground for two important facilities: 1) the Administration, Research & Engineering Center in Sunnyvale, CA and 2) the Development & Test Center in the nearby cattle grazing lands in the foothills south of San Jose, CA, near the community of Coyote, CA.

CSD’s Facility in Sunnyvale. CSD moved into this facility in October, 1960

2

CSD’s Development & Test Facility in Coyote Valley (about 5,400 acres), also serving as a

game reserve for elk, deer, mountain lions, bobcats, coyotes, wild boar and many birds In June of 1960, NASA awarded CSD a contract to test a 1-segment motor (TM-4). The purpose of this test was to prove the feasibility of segmentation for reasonably large solid rocket motors. The TM-4 had a 42-inch diameter and an overall length of about 7-ft. The forward and aft closures were attached to the segment using two joint configurations: 1) a bolted flange and 2) a clevis joint.

TM-4, NASA 1-Segment Motor, Tested 12/15/1960, Thrust = 15,000-lb With the success of the TM-4 test, CSD selected the clevis joint for further development. CSD quickly moved forward to appreciably increase the size of segmented test motors with their design and test of the P-1 motor for NASA and the P-1-2 motor for the Air Force. The P-1 was also a 1-segment motor; however, the diameter was increased to 90-in. The segment was conical-shaped with a conical-shaped propellant bore. The intent of the conical bore was to increase the port cross-sectional area as the accumulating gases from the burning propellant flowed toward the nozzle, thus maintaining a near-constant Mach number down the bore. This motor had an overall length of 26-ft and a firing duration of about 80-sec. The P-1-2 had the same forward segment, but added a second segment at the aft end with a diameter of about 100-in. This motor had an overall length of about 40-ft and a duration of about 80-sec. Both tests were highly successful and were a critical element in providing the Air Force

3

with the confidence that very large strap-on solid rocket boosters could be used in conjunction with a Titan-II liquid propellant core. It was soon determined that conical-shaped segments were not necessary and could be replaced with a cylindrical configuration, thus providing simplicity of manufacturing and interchangeability of segments for such programs as the Titan SRMs. In early 1962, a P-1-2 was static tested, for the Air Force, with a duration of 130-sec in order to demonstrate nozzle material survivability.

P-1 (1-segment motor) P-1-2 (2-segment motor) Tested 8/61, F = 220,000-lb Tested 12/61, F = 500,000-lb Titan Boosters During early 1962, after a tough competition, CSD was awarded a significant Air Force contract to develop 5-segment, 10-ft diameter, strap-on boosters with 425,000 lb of PBAN propellant for what was entitled the Titan IIIC Program (Air Force designation: 624A). One must remember that this was back in the days when mainframe computers were in their infancy and PCs/laptops did not exist. Comprehensive computer programs for accurate predictions of material and motor performance did not exist. This was the time when slide rules and Frieden calculators were the mainstay on the engineer’s desk. Propellant grain designers used drawings from the drafting department on which were depicted burnback estimates for each grain configuration. Then, with the use of a planimeter and a map reader, the engineer estimated burn surface areas for each increment of web burnback—a slow and tedious process. This was a time during which much testing was conducted to demonstrate the technologies and designs being developed. These new strap-on boosters were going to be many times larger than any previous SRM which had ever flown. Burn times over 100-seconds created huge demands for motor case insulation systems and nozzle liners. Liquid injection thrust vector control (LITVC), in the exit cone of the nozzle, was selected to turn the direction of the booster’s thrust vector, when steering was needed. Choosing a reactive liquid, N2O4, had the benefit that the injectant added to the overall specific impulse (Isp) of the SRM. Controlling the relative thrust between two very large boosters, in order to maintain vehicle flight stability, required close attention to propellant burn rate reproducibility, grain design details and the resultant thrust histories. Ignition of the two large boosters had to be controlled to within milliseconds (msec) so as to minimize delta thrust between the boosters. A very large pyrogen igniter (with its own smaller pyrogen igniter) was used to ignite each booster within 250 msec, allowing the two boosters to ignite within 20 msec of each other. This was indeed a pioneering effort that would pave the way for other solid rocket strap-on boosters for many decades to come. Another important design feature which added to vehicle flight stability was the slight outward cant (6º) of the SRM nozzles. CFD modeling did not exist to help define pressure distributions and Mach numbers down the propellant grain. A single time step calculation, on the mainframe, for flow conditions down the

4

grain, took 45 minutes with only approximations being provided. In order to better understand the critical design elements, CSD conducted many subscale motor firings which simulated the full-scale motor configuration. Additionally, many cold flow tests were conducted. The cold flow test equipment modeled the segmentation of the SRM propellant grain and had: 1) flow inlets along each segment to help account for mass addition along the grain and 2) pressure transducers along each segment and within the slots (between the segments) to determine pressure distributions along the propellant grain. This evaluation was extremely important in order to: 1) understand the variation in propellant burn rate along the grain (due to a pressure drop within the motor), 2) establish sizing parameters for the thrust termination system for the early Titan IIIC SRMs and 3) use pressure distribution data (along the bore and within the slots) in conjunction with propellant structural properties to prevent catastrophic grain deformations (aft of slots) which might form a “propellant nozzle” forward of the actual nozzle, possibly causing the motor to over pressurize and rupture the motor case. For the Titan Program, CSD was a total booster stage contractor (for Stage Zero) with complete responsibility for:

• The large Solid Rocket Motors (SRMs), which were the primary part of Stage Zero • All Stage Zero structures, including support structure for the liquid propellant core, aft

skirt, nosecone and the side-mounted TVC tanks (originally produced at Sikorsky, later at CSD)

• Separation SRMs (4 forward and 4 aft on each booster) to pull the two boosters away from the liquid propulsion core, near booster burnout

• All Stage Zero electronics, including power supplies, and Ground & Flight Instrumentation (FI)

• All Destruct and Ignition Safe & Arm (S&A) devices including Destruct Ordnance and an Inadvertent Separation Destruct System (ISDS)

• Stage Zero thrust vectoring system: Liquid Injection Thrust Vector Control (LITVC), using 24 valves per nozzle to determine flow quantity and location for the injection of N204 into the nozzle flow stream---for the Titan IIIC series the valves were controlled hydraulically, for Titan 34D and Titan IVA configurations, electromechanical valves (EMVs) were used. Moveable nozzles were examined for the Titan IIIC SRM application, but were not considered mature enough for very large SRMs in the early 1960s.

• On early flights, CSD was also responsible for a booster Thrust Termination (TT) system which involved openings through each SRM’s forward closure propellant grain, explosively released port covers on the forward dome and flow channels within the nosecone to direct the negative thrust.

Loaded Titan Segment Removed from an In-ground Oven

5

Titan booster build-up and aerial view of static test at CSD’s ST-9 vertical test stand Besides the nozzle-up static testing of the Titan SRMs at CSD’s facility near San Jose, CA, some special testing was conducted at Edwards Air Force Base (EAFB) in Lancaster, CA and the Naval Ordnance Test Station (NOTS) in China Lake, CA. Since SRMs of this size had never previously been tested, it was important to determine the safety of such systems, as well as one’s ability to command the destruction of a large SRM, if needed. Additionally, the Titan SRM Thrust Termination System also needed to be proven. The remoteness of EAFB and NOTS was ideal for such testing. To demonstrate the safety of a large SRM, a sled test was devised at NOTS in which one Titan SRM segment motor was placed on rails and slammed against a large cement wall to prove that detonation did not occur upon impact. In a test at EAFB, a 5-segment SRM was strapped in a horizontal orientation, ignited, followed by thrust termination, and then successfully commanded to destruct. These tests were conducted by Air Force personnel working in conjunction with CSD personnel stationed at that facility. Additionally, the Air force built a vertical nozzle-up test stand at EAFB as a backup to CSD’s ST-9 facility. CSD tested two development and three qualification Titan SRMs in the EAFB stand. Concurrent use of both CSD’s and EAFB’s static test stands appreciably shortened the schedule to the first flight. CSD took the 5-segment booster from drawing board to its inaugural flight in less than 3-1/2 years. The 5-segment version of the Titan booster had a total of 35 flights with 100% booster success. Within these 35 flights there were three vehicle configurations: Titan IIIC, Titan IIID and Titan IIIE, the differences being in the Titan core and payload fairing. Titan IIIE, specifically, had the important change of including Pratt & Whitney’s cryogenic liquid engine, the RL-10, which was part of the Centaur Upper Stage. Payloads included Air Force satellites as well as NASA scientific missions including Viking 2 & 1, which were launched in that order toward Mars in 1975 (landing in 1976), and Voyager 2 & 1 which began their grand tour of the planets and then their marathon exploration of our galaxy in 1977. Many of the Titan III configurations carried multiple payloads on a single flight. In late 1967, the Air Force Manned Orbiting Laboratory (MOL) Program required the Titan boosters be increased in length by adding two additional segments (7 total) and by increasing the length of the forward closure (to provide the desired thrust history regressivity). CSD’s ST-9 test stand had to be modified with a steel superstructure to accept the additional length of the 7-segment motor. The MOL Program was canceled in June 1969 by which time the first 7-segment SRM had been tested. The SRM test program, however, was continued until three additional static firings were completed in 1970. These initial four 7-segment static firings were critically important because they qualified electromechanical valves (EMVs) for the LITVC system. Also qualified was an ullage blowdown TVC system. This qualification allowed the replacement of the hydraulic valves with EMVs on the remaining 5-segment SRM flights (Titan IIIC/D/E). Both of

6

these features were also incorporated into the Titan 34D SRMs discussed below. The Air Force/NASA would return to a 7-segment Titan booster in the late 1980s, at which time two additional static firings were conducted (late 1987 and Valentine’s Day 1988) to complete its qualification for flight. In the interim period between the flights of the 5-segment and 7-segment Titan SRM configurations, the Air Force had CSD add a “half” segment for additional flight performance over the Titan IIIC. This new configuration was called the Titan 34D. Its inaugural flight was a night launch on 30 October 1982. Besides the Air Force & NASA use of the Titan 34D, Martin Marietta (now Lockheed Martin) used the same booster configuration for a commercial version of Titan. Four of these Commercial Titan (CT-III) vehicles were launched. The loss of Shuttle’s Challenger in 1986 caused the Air Force to make sure they had independent access to space for their large payloads. To meet the increasing demand for a higher payload capability, CSD was directed to complete the qualification of the 7-segment Titan booster. This Titan vehicle, the Titan IVA, had its inaugural launch in June 1989.

Titan III Titan 34D Titan IVA 5-Segment SRM 5-1/2-Segment SRM 7-Segment SRM First Launch: 6/18/1965 First Launch: 10/30/1982 First Launch: 6/14/1989 Wp = 425,000-lb/SRM Wp = 464,000-lb/SRM Wp = 591,000-lb/SRM FVac = 1,310,000-lb/SRM FVac = 1,400,000-lb/SRM FVac = 1,600,000-lb/SRM Burn time = 115-sec Burn time = 114-sec Burn time = 120-sec Titan launch vehicles (T-III Series, T-34D and T-IVA) were used for a wide variety of scientific, communications and Air Force payloads, some of which are shown below:

Mars Viking Lander 2 & 1 Voyager 2 & 1 Mars Observer Air Force Payloads 1975 1977 1992 1965-1997

7

Each of the Titan SRMs used a similar case segment/grain design. The forward closure contained a star grain to help provide the needed thrust regressivity. The segments (120-in diameter and about 120-in long, each with 78,000-lb of propellant) used a conical-shaped bore to produce a tailoff of about 10-sec. This tailoff was needed to control the thrust differential between the two strap-on boosters when nearing booster burnout and separation. The forward end of each segment grain was inhibited (using an insulation restrictor material) from burning in order to help produce the required thrust history. Each segment-to-segment and segment-to-closure interface used a clevis joint, fastened with a series of clevis pins positioned circumferentially around each joint. The clevis interface was oriented such that the male portion pointed upward with the female portion pointing downward. This was done to minimize the possible ingestion of water/moisture. A single o-ring (made with a mold which produced a continuous product with no splices) was used at each interface. A total of 153 such o-rings were used in static tests and 1,398 of these o-rings used in flight. CSD never experienced an o-ring failure.

Typical Titan Clevis Joint, Showing the propellant-insulation-case interfaces A total of 106 Titans (using CSD’s Titan III, 34D and IVA booster configurations) were launched. Of these, there were two flight failures caused by the Titan SRMs: 1) a Titan 34D and 2) a Titan IVA. On 18 April 1986, a Titan 34D was launched from the Vandenberg launch facility in California. Shortly after liftoff, the vehicle exploded. A thorough investigation established that Segment #1 (lowest position segment) on one of the SRMs experienced overheating and eroding of the motor case near the forward joint of the segment, causing it to fracture, resulting in the destruction of the vehicle. The failure analysis showed that a faulty bond had occurred between the case and insulation. The poor bond resulted from insufficient controls within the 20 year old manufacturing methods used at the case/insulation supplier and the absence of adequate Non-Destructive Evaluation (NDE) of the segments. On 2 August 1993, a Titan IVA was launched from Vandenberg. Late in the SRM burn, a case burnthrough occurred, causing the liquid core to explode. The subsequent investigation showed the problem was caused by a misapplication of a procedure to repair the grain restrictor at the

8

forward end of Segment #3 (3rd from the bottom) on one of the SRMs. Statistical analysis of segment processing results, NDE data and processing parameters resulted in a new set of process limits to control generation of defects. It is important to note that neither Titan SRM failure was due to a design shortcoming. It does clearly show that great attention needs to be given to each manufacturing and processing detail. The Titan SRM Program revolutionized the SRM industry with its segmentation. This segmentation allowed an order of magnitude increase in propellant weight over previous SRMs, while still allowing rail and road transportability. The Titan SRM program established an important experience base which greatly benefited other programs using segmented solid rocket boosters (shown below):

Space Shuttle Titan IVB Japan’s H2 Europe’s Ariane 5 Titan Launch Support Teams CSD had complete stage (Stage Zero) responsibility for the Titan III boosters. Therefore, CSD provided launch services at both the Cape Canaveral and Vandenberg launch complexes. CSD personnel from Coyote and EAFB provided the nucleus of these teams because of their extensive experience with the Titan SRMs. These teams, of approximately 150 personnel at each facility, had the responsibility of assembling the boosters for launch, completing all checkout operations and supporting the Air Force and Martin Marietta teams in the launch of the Titan vehicles.

Aerial View of Titan III Launch Complex at Cape Canaveral Air Force Station (CCAFS) All CSD segments and closures were shipped to the launch sites via railroad flat cars. The convenient segment length (the result of Aerospace Ground Equipment [AGE] studies) of about 10-ft allowed each segment to be shipped vertically in air conditioned shipping containers. Upon receipt/inspection at CCAFS, the SRMs were assembled in the Solid Motor Assembly Building (SMAB). The core was integrated in the Vertical Integration Building (VIB). The vehicle was mated on a transporter railcar which included a mobile launch tower. The entire launch system was then transported to either of two Titan launch pads. The two Titan launch pads allowed

9

simultaneous and independent processing of two Titan Launch Vehicles. This experience base provided important insight for some of the launch site operations used for the Space Shuttle. At Vandenberg, the SRMs were assembled at the launch pad.

Titan Segments Shipped by Solid Motor Assembly Bldg. SRM Assembly Railroad Rail to both Launch Sites (SMAB) at CCAFS Car, Including Mobile Service Tower (MST) at CCAFS Algol III In 1971, CSD was awarded a contract to design, test and manufacture a new booster stage for NASA’s Scout Launch Vehicle. CSD’s motor, the Algol III, used a 45-in diameter monolithic steel motor case which contained about 28,000-lb of propellant. The cross-sectional loading of the propellant was very high, producing a small port-to-throat ratio, resulting in an operational condition that would create high erosive burning. To combat this situation, CSD coated the forward portion of the grain’s mandrel with a wash coat of a variant of the motor’s PBAN liner. The coating bonded to the grain, slowing ignition and nicely solving the problem. Algol III had a fixed ablative nozzle. Steering was provided by customer-supplied jet vanes which were located in the nozzle flow stream, aft of the nozzle. These vanes were directly attached to the movable aerodynamic fins on the vehicle. Algol III was used for 35 Scout Vehicle launches, and performed with 100% success.

Algol III Booster Launch of a NASA Scout Vehicle

10

Separation Motors for Large Solid Boosters ----Titan: CSD was responsible for the design, test and manufacture of the rocket motors used to separate the Titan solid rocket booster from the Titan III liquid propulsion core. Each booster had four of these separation motors at each end for a total of sixteen per launch. The Titan Staging Motors had a monolithic aluminum case with a 6-in diameter, had a slightly canted ablative nozzle, were about 5-ft long and contained approximately 55-lb of PBAN propellant. Each motor produced about 4,500-lb of thrust. These motors were a direct derivative of the CSD’s Titan II Translation Motors which were also used for some Martin Marietta retro rocket demonstrations. The Titan SRM staging motors (1,648) were 100% successful in flight.

Titan Staging Motor being Hoisted to its Titan Nosecone Location ----Shuttle: In 1975 CSD was awarded a NASA contract to design, test and manufacture the Booster Separation Motors (BSMs) for the Space Shuttle Solid Rocket Strap-on Boosters. Like the Titan III strap-on boosters, the Shuttle required four BSMs at each end of each strap-on. Since the BSMs have to operate in close proximity to the Shuttle Orbiter, a very stringent requirement was imposed on the operational characteristics of these motors- NO EJECTA (for BSMs operating forward of the Shuttle Orbiter, i.e., those located at the forward end of the solid rocket boosters). The “no ejecta” requirement resulted in some unique design characteristics for the CSD BSMs:

1. A very low level of aluminum (2%) in the propellant 2. A nozzle with a monolithic graphite throat and a steel exit cone located aft of the throat

where no melting occurs 3. A nozzle environment closure (for forward located BSMs only) that remains with the BSM

after opening. These forward-mounted nozzle closures were made of steel, had a strong-back for stiffness, were hinged on one side to the nozzle exit cone and were held closed by a frangible link which ruptured upon BSM ignition. Internal mechanisms, within the closure design, prevented the closure from hitting the Shuttle SRB nosecone and from bouncing back into the flow stream of the BSM nozzle. These forward BSM nozzle closures were known as “Aeroheat Shields” (AHSs) because they protected the forward BSMs from the potential of autoignition caused by aeroheating during the ascent of the Shuttle after launch. The aft-mounted BSMs (located aft of the Shuttle Orbiter) simply used a sealed disk closure which detached and blew free of the BSM upon ignition.

11

The CSD BSMs had a 12.8-in diameter aluminum case which contained 72.9-lb of an 86% solids hydroxy-terminated poly-butadiene (HTPB) propellant. The 18-point star grain helped produce the needed 22,000-lf of thrust. With the nozzle canted at 20º, the motor had an overall length of 34.6-in, including the through-bulkhead initiators (TBIs) of the ignition system. Over 1900 CSD BSMs were used for Shuttle flights with 100% success.

BSM Cross-section & Nozzle Cover

4 Forward BSMs 4 Aft BSMs Formation of a New Propulsion Integration Company—USBI Capitalizing on its Titan booster integration experience, CSD, in 1977, formed a subsidiary called United Space Boosters Incorporated (USBI), under the direction of Dr. Frank Lavacot who was the former Director of CSD’s Development and Test Center. Dr. Lavacot assembled this new organization using key CSD personnel with extensive booster manufacturing and integration experience. The team’s unique heritage allowed USBI to compete and win the Shuttle booster integration contract. Initially, USBI was co-located at both Huntsville, AL, and Kennedy Space Center (KSC) in Florida. USBI was responsible for facility activation, Shuttle Solid Rocket Booster (SRB) assembly & integration, launch support, range safety system, recovery and disassembly. SRB integration involved responsibility for all SRB structures beyond the SRMs, including avionics, separation systems and parachute systems, etc. When facilities became available, all operations were moved to KSC.

12

USBI also oversaw the building and outfitting of two SRB recovery ships. Additionally, USBI refurbished and re-certified SRB hardware after recovery and return of the SRBs to the CCAFS docks. As part of this refurbishment process, USBI, in conjunction with MSFC, developed a spray device which was used to remove post-flight insulation from the exterior of reusable SRB components (i.e., aft skirt, etc) and then used the spray system to apply a new layer of insulation. Variants of this spray system eventually had important commercial applications. Many years later, USBI was integrated into the United Space Alliance (USA) Space Shuttle Team.

Dr. Frank Lavacot Shuttle Booster Recovery Ship

Items in Color (Red Arrows) Indicate the Shuttle SRB Items with USBI Responsibility

Upper Stages and Space Motors

CSD upper stages and space motors used filament wound motor cases in order to provide high mass fractions for their systems. Early motors cases were made of 901S fiberglass-epoxy. Later the material of choice became Kevlar-epoxy with its higher strength and modulus qualities. Eventually newer systems were made with graphite-epoxy motor cases.

13

All of the filament wound motor cases stretch under pressurization. CSD took advantage of this trait by pioneering a process which involved curing these motors with internal pressure applied to the motor case during the propellant curing process. This “pressure cure” helped make up for most of the shrinkage which normally takes place during the propellant curing process. The result was a near “stress free” condition for the propellant system while in storage. This appreciably improved the propellant aging characteristics within the motor. An additional characteristic of CSD’s space motors was that almost all grains were machined on the internal surfaces to their final dimensions. This provided two important benefits: 1) the total impulse of the motor was accurately controlled, improving reproducibility and 2) it controlled the burnout characteristics of the grain, virtually eliminating tailoff when desired. Upper stage and space motors inherently have very high nozzle expansion ratios. Therefore, testing such motors in test facilities which have near-sea level ambient pressures creates thrust and Isp values which are not representative of operation in space. Additionally, for motors using very lightweight exit cone materials, high expansion ratio nozzles could implode when tested near sea level. In order to remedy these concerns, CSD tested this class of motors at Arnold Engineering Development Center (AEDC), located in south-eastern Tennessee. AEDC was the premier US facility providing near-vacuum ambient test conditions for upper stage motors. This facility is renowned for its very professional handling of their customer’s motors and for the accuracy of their measurements of each motor’s performance. FW-4 In 1964 CSD won a NASA contract to build an upper stage/kick stage. The resultant motor used a filament wound fiberglass-epoxy motor case of 19.6-in diameter and contained 605-lb of PBAN propellant. This motor, named FW-4, was used as an upper stage for the Scout Launch Vehicle as well as a kick stage for the Thor, Atlas and Delta II launch vehicles. On some missions, the FW-4 was spun up to 200 rpm during its operation. There were 67 flights of the FW-4, with only one failure. This failure was associated with a nozzle component flaw which was pointed out to NASA before the flight. NASA decided to accept the risk and proceeded with the launch.

FW-4 Filament Wound Fiberglass-Epoxy Motor FW-5 In September, 1970, TELSAT Canada awarded Hughes Aircraft Company the contract to build Canada’s first communications satellite named “Anik”. The Anik-A was the world's first national domestic satellite. That same year, CSD was awarded the contract to build the apogee kick motor for the mission. The motor, designated FW-5 by CSD (HS-333 by Hughes), had a 28.6-in diameter filament wound fiberglass-epoxy case, a length of 44-in and contained propellant loads from 518 to 598-lb, depending on the mission. An interesting test requirement during FW-5’s development was to pre-cool the ablative nozzle exit cone to -200°F. Although the case and nozzle technology came from the proven FW-4 program, the FW-5 propellant used an HTPB

14

binder system. Besides the Anik missions, FW-5 provided the apogee kick for other missions such as Palapa, Westar and other satellites. In all, FW-5 flew 17 times with 100% success.

FW-5 Filament Wound Motor Anik-1 Satellite Inertial Upper Stage (IUS) In the early 1970s, the Air Force and NASA were looking at a new space tug which they called Interim Upper Stage (IUS). This system had a baseline approach which used liquid propulsion systems. At this time, NASA thought they would charge Shuttle Payload Bay users based on either of two criteria: 1) the percent of total Shuttle Payload Bay length occupied by the spacecraft/payload vehicle or 2) the percent of total Shuttle payload weight capacity consumed by the combination of spacecraft/payload. In 1973, CSD’s Dr. Altman thought that NASA should take advantage of the density-impulse advantages of solids versus liquids. Extensive trade studies were conducted at CSD. The results convinced Boeing that this was the approach to take, and a team was formed. The Boeing/CSD team then competed against the various liquid approach teams and won. This then created a new competition amongst various SRM manufacturer/prime contractor teams that were formed. The Boeing/CSD team won. In early 1976, the Boeing/CSD team was awarded an 18-month Validation Phase Program which required a full-scale static firing demonstrating the technologies required for this application of solid propulsion systems. In December 1977 (only 15 months after go-ahead), the test was successfully conducted at Arnold Engineering Development Center (AEDC). Having the needed confidence to proceed, the Air Force awarded the IUS Full-scale Development Program (FSD) to the Boeing/CSD team in early 1978. At some point during this FSD program, the “I” in IUS was changed from “Interim” to “Inertial”. Again, CSD was a “stage” contractor with full responsibility for all propulsion, thrust vectoring and electronics for the resulting 2-stage (Orbus 21 1st stage and Orbus 6 2nd stage) propulsion system. Early in the program, the Air Force decided it needed more payload capability and added the requirement that the Orbus-6 add an extendible exit cone (EEC). This version of the 2nd stage was named “Orbus 6E”. With the addition of the EEC, the nozzle expansion ratio was increased from 49.3:1 to 181.1:1. This increased the delivered Isp of the motor by 14-sec. In the stowed condition, the EEC only added 4.2-in to the motor’s overall length. The Air Force’s confidence to incorporate an EEC

15

came from pioneering work done by CSD during Independent Research & Development (IR&D) programs at CSD.

Extendible exit cone(optional)

Thrust vectorcontrol actuator

Nozzle

First stagelid rocket moto

so r(O

Interstage

Reaction control system

ketE

Spacecraftseparation plane

Environmentalmeasuring unit

(EMU)

Avionics bay (redundantinertial measurement

unit — RIMU)

rbus 21)

Second stage solid rocmotor (Orbus 6 )

1st Stage Orbus 21

2nd Stage: Orbus 6E

Air Force/NASA IUS, built by Boeing, a 2-Stage Space Vehicle using CSD’s Orbus 21 and

Orbus 6E Solid Propellant Rockets. It was Configured to Fly off both the Shuttle and Titan Launch Vehicles

Orbus 21: IUS 1st Stage Orbus 6E: IUS 2nd Stage Diameter = 92-in Diameter = 63-in Wp = 21,400-lb Wp = 6,000-lb

16

otors were responsible for a number of important milestones in the solid propulsion indu

1. 1 was the largest solid rocket space motor. Diameter of 92-in, overall length =

2. e he use of

3. m (Techroll

4. dundantly driven electromechanical

5. hest delivered specific impulse (304.3 sec.) of any operational

6. was

ns in order to maintain the same

7. ere some of the first flight systems using filament wound Kevlar-epoxy

8. ad to withstand the highest line loads of any composite motor case to-

date (3,560 lb/in)

ttle,

as

the Shuttle. After the huttle maneuvered to a safe distance, the IUS first stage was ignited.

The IUS m

stry: Orbus 2124-in Orbus 21 had the longest burn time (145-sec at 60ºF) of any solid motor, including thShuttle boosters. This created significant nozzle challenges, resulting in tcarbon-carbon for both the nozzle throat and exit cones of both motors. Both stages incorporated the world’s lowest torque movable nozzle systeSeal). This system helped minimize the weight of the nozzle structure. Both stages were the first to incorporate a reactuation system for vectoring the nozzles. Orbus 6E produced the higsolid rocket flight system. Both stages had to be qualified for propellant loads between 50% and 100%. Offloads were machined as a “burnback” of the original grain. Additionally, the nozzle throat machined (“pre-eroded”) for offloaded configuratioMaximum Expected Operating Pressure (MEOP) Both stages wmotor cases. The Orbus 21 was a key structural member of the IUS vehicle and as a consequence the motor case skirts h

The Boeing/CSD IUS vehicle was configured for launch on both the Titan and Space Shuttle launch vehicles. On the Titan vehicle, the IUS was attached via an interstage structure atop the upper stage of the liquid propulsion core. On the Shuttle, the IUS was carried in the “horizontal” orientation within the Shuttle Payload Bay. Therefore, before ejection of the IUS from the Shuit was necessary to elevate the forward end of the IUS 58° above the longitudinal axis of the payload bay. The aft end of the IUS was attached to a “Tilt Table”. Rotation of the Tilt Table waccomplished using a variant of the actuators CSD designed for vectoring the nozzles on bothIUS SRMs. A spring loaded frangible ring then separated the IUS from S

IUS Tilt Table CSD Actuator

Oriented for IUS ejection Shuttle Payload Bay: IUS Ejection Shuttle with IUS Tilt Table The maiden flight of IUS took place on 30 October 1982 aboard the maiden flight of a Titan launch vehicle using CSD’s Titan 34D boosters. This launch placed a DSCS-II and a DSCS-III

17

satellite into what was termed “bull’s eye” orbits. A total of 24 IUS vehicles were launched over about a 20 year period. On two of the missions, the IUS vehicle did not have a chance to perform: 1) Challenger and 2) a Titan IVA on which the payload fairing did not properly open/separate. missions included such payloads as the initial NASA Tracking & Data Relay Satellite (TDRS) communications constellation, many Air Force satellites, Magellan Spacecraft to map the surface of Venus, Galileo Spacecraft to perform years of experiments around Jupiter, Ulysses Spacecrafto orbit the Sun and the X-ray Telescope—Chandra. All Orbus 21 motors performed flahowever, on the TDRS-1 mission (April 1983) the Orbus 6 had a nozzle vector system overheating anomaly. This resulted in design changes to improve the local insulation system near the movable joint of the nozzle. The TDRS-1payload, ho

IUS

t wlessly,

wever, still made it to its intended rbit. TDRS-1 was recently retired after 26 years of service.

o

TDRS Magellan Galileo Ulysses

Some of the IUS Spacecraft

) to

Orbus

21S

) which released itself upon motor ignition. All flights of the rbus 21S were 100% successful.

ttle ts

eotransfer responsibility. The mission achieved its original geosynchronous objective.

Orbus-21S for Intelsat VI The Hughes Space and Communications Group’s Intelsat VI (I-VI) series of satellites (1983-1991were the largest in the world at the time. In the early 1980s, Hughes awarded CSD a contract provide the perigee kick stage. This motor (Orbus 21S) was a spin stabilized stage requiringqualification with spin rates up to 36 rpm. The Orbus 21S required added insulation over its Orbus 21 predecessor because of the spin environment. The propellant formulation for the 21S was identical to the Orbus 21 of the IUS program. The propellant load had to have an offload capability ranging from 0 to 50%. This was accomplished by machining the propellant bore (to a configuration similar to the normal burnback patterns of the grain) until the desired total propellant weight was achieved. Unlike the Orbus 21 for IUS, all nozzle throats of the Orbus were machined for the offloaded configuration. The customer opted for the slight Isp penalty resulting from the lower nozzle expansion ratio, because nozzle throat logistics were simplified. The Techroll Seal of the IUS was replaced with a solid structure since no TVC was needed for theOrbus 21S. Another special design feature was a lateral support system for the nozzle exit cone(to handle flight loads during launchO During a March 1990 launch on a Commercial Titan, the separation of the I-VI spacecraft from the upper (second) stage of the Titan failed. This prevented the Orbus 21 S from carrying thesatellite to a higher orbit. The control room commanded the Intelsat VI to separate from the Orbus 21S and, using the satellite’s on-board propulsion system, raised the satellite to a low-earth parking orbit. In May 1992, a Space Shuttle carrying a new (replacement) Orbus 21S rendezvoused with the I-VI. Three astronauts captured the satellite and attached it to the new Orbus 21S. After ejecting the Intelsat-VI/Orbus 21S assembly from the payload bay, the Shumaneuvered to a safe distance and the Orbus 21S was ignited to successfully carry out ig

18

Orbus 21S used to Orbus 21S and Intelsat VI Rescue Intelsat VI Leaving the Shuttle Bay Transfer Orbit Stage (TOS) Orbital Sciences Corporation worked with NASA and Martin Marietta to provide another space vehicle which was simply called Transfer Orbit Stage (TOS). The transfer rocket chosen for this application was the Orbus 21, already proven on IUS missions. Two flights of TOS took place: 1) on 25 Sept 1992 Mars Observer was launched by the 4th Commercial Titan (a payload systems problem occurred when the payload reached Mars) and 2) in September 1993, the Advanced Communications Technology Satellite (ACTS) was launched by the Space Shuttle and was successfully placed in earth orbit. On both missions the Orbus 21 performed flawlessly.

Mars Observer ACTS Athena Launch Vehicle Upper Stage: Orbus 21D A partnership was formed in the mid 1990s among Lockheed Martin, Thiokol and CSD. This partnership involved making available small launch vehicles called Athena I and Athena II. The Athena I used a Castor 120 first stage and a CSD Orbus 21D second stage. For the Athena II, both the first and second stages were Castor 120s. The third stage was an Orbus 21D. The Orbus 21D was basically an Orbus 21 with the nozzle exit cone changed from a lightweight carbon-carbon system to a robust ablative system using carbon phenolic. The added structural qualities of the phenolic were needed for “hot” separations where the upper stage was ignited in close proximity to the previous stage. Seven missions were launched. Two flight failures occurred, but none attributable to the Orbus 21D which performed with 100% success. Successful missions included NASA’s Lewis Satellite. Lunar Prospector, Rocsat-1, IKONOS-II and NASA’s Kodiak Star.

19

Orbus 21D Stacking Orbus 21D Athena-1 Athena-II with Static Test on Athena-I Launch Lunar Prospector

CSD’s First Complete Launch System Kangaroo Meteorological Rocket From 1968 into the early 1970s, CSD was involved in a small project that gave CSD their first experience in working with a complete vehicle, albeit only 6-1/2 inches in diameter. This solid rocket vehicle was the Kangaroo Meteorological Rocket, the mission of which was to launch a meteorological dart to an altitude of 400,000-ft. A unique feature of the Kangaroo was that the dart (usually mounted atop the booster) was housed within a canister inside the motor’s combustion chamber. An end-burner grain at the base of the canister determined when motor chamber pressure reached the dart’s aft end. At the correct balance of pressure loads and flight acceleration loads, the dart was ejected from the canister. Moving forward the dart broke open the vehicle’s frangible nosecone, sliding through the dart’s fins housed within the nosecone. The aft end of the dart was flared to pick up the fin assembly when exiting the canister. The booster burned out at 18,000-ft, but the aerodynamic efficiency of the dart carried the dart to 400,000-ft. An additional requirement was that the Kangaroo booster have stable flight after dart separation and return to the earth in a predictable manner. This resulted in the vehicle having unusually large fins. Two Kangaroo’s were launched from San Nicolas Island, off the coast of Pt. Mugu, CA. Radar did not successfully track the birds. A subsequent launch at White Sands, NM was fully successful. Unfortunately, shortly afterward, the Navy decided they did not need the system.

Prelaunch Checkout of Kangaroo White Sands Launch

20

Missile Propulsion Systems Mk-111 Tomahawk Booster In the early 1980s, CSD won the Navy’s Mk-111 booster program for the Tomahawk missile. A key element of the win was the successful demonstration tests conducted during the competition. These tests included advanced technologies which helped the Mk-111 achieve its performance goals. This program was particularly important to CSD because it began a long and very successful history for CSD in the tactical and missile defense field. The Mk-111Booster requirements were significantly increased over the previous Mk-106 Booster. The total delivered impulse was significantly increased, the storage/operating temperature range broadened, missile and launcher interfaces maintained while the motor envelope and weight were unchanged from the Mk-106 Booster. The Mk-111 Booster incorporated the following key design features:

• A high performance high solids HTPB propellant • A very high volumetric loaded propellant grain design • An efficient Ball and Socket Movable Nozzle Thrust Vector Control System • A sophisticated volume efficient Hydraulic Nozzle Actuation System

Of significant importance was CSD’s incorporation of the Ball and Socket movable joint for the Mk-111 nozzle. This system was subsequently used on several other CSD tactical motor programs including the Mk-72 Booster and the Theater High Altitude Area Defense (THAAD) missile booster. CSD’s Mk-111 provided superb flight reliability (100% success).

Mk-111 Booster Tomahawk Launched from a Submarine using CSD’s Mk-111 Booster Trident II (D5) In October 1983, CSD was awarded a contract to design, test and manufacture the 3rd stage of the submarine-launched Trident II (D5) missile. CSD’s 3rd stage motor development program was the most successful large motor development program ever conducted in the solid rocket industry, with no motor or component failures in any of the static or flight tests of the development program. The 3rd stage also delivered 100% success on all subsequent flight tests.

21

The 3rd stage originally used a filament wound Kevlar-epoxy motor case which was subsequently upgraded to graphite-epoxy in order to reduce motor inert weight and thus improve flight performance. The propellant system was a polyethylene glycol, nitroglycerin-plasticized (PEG/NG) system, with a Department of Transportation (DOT) hazard class rating of 1.1.

Trident II (D5) submarine launch with CSD’s 3rd stage propulsion aboard Mk-72 Standard Missile Booster In the mid-1980s the US Navy, in conjunction with Raytheon/Standard Missile Company, initiated an effort to upgrade the fleet air defense system by upgrading the Standard Missile. The improved missile incorporated a booster as a first stage of a two stage missile. The new missile would increase missile range and intercept altitude as well as incorporating high initial maneuverability immediately after launch to intercept incoming cruise missiles at low altitude. The high maneuverability could not be achieved using aerodynamic fins because the missile was at insufficient velocity immediately after launch. Therefore, the new booster needed a thrust vector control (TVC) system to enable the high pitch over maneuver. Volumetric constraints were also imposed to allow this missile to be contained in the Aegis ship Mk-41 Vertical Launch System (VLS). CSD won this program utilizing a four-nozzle Ball and Socket TVC system which incorporated an electro-mechanical actuation (EMA) system. CSD clearly demonstrated, in the proposal phase, the technologies needed to minimize program risks. Both a two-nozzle and a four-nozzle prototype motor configuration were successfully static tested. The booster incorporated a very high solids loaded HTPB propellant. The motor went through a comprehensive development and qualification program and became operational. The Mk-72 booster is utilized on several versions of SM-2 and SM-3 missiles and has demonstrated 100% success in operational flights.

Mk-72 Booster Stage SM-3, Block II Shipboard Launch

22

Orbus 1 Upper Stage/NMD-GBI Orbus 1A The Orbus 1 Upper Stage was developed and qualified in the late 1980s as the 3rd stage of the Strategic Target System (STARS) vehicle for the Ballistic Missile Defense Organization (BMDO), of the Department of Defense, through Sandia National Laboratories. It has also been used as the 3rd and 4th stages of the Starbird vehicle. More recently (late 1990s), the Orbus 1 was selected as the 2nd and 3rd stage (Orbus 1A) of the National Missile Defense/Ground-Based Interceptor (NMD/GBI) missile through McDonnell Douglas/Boeing, but was later terminated after it was determined that the GBI missile required more energy and larger missile/motors. The Orbus 1 motor used the latest state-of-art technologies packaged to meet envelope and performance requirements. It incorporated a graphite epoxy case, 90% solids HTPB propellant with HMX, a flexseal movable nozzle, and an electromechanical actuation system. The motor diameter was 27.6 inches with an overall length of 49.6 inches.

Orbus 1 Stages 2 & 3 for the Starbird Launch Vehicle Terminal High Altitude Area Defense (THAAD) In the early 1990s CSD was selected by Lockheed Martin to be its teammate to develop and qualify the Terminal High Altitude Area Defense (THAAD) propulsion system. Similar to the Mk-72, requiring high maneuverability, a high performance Thrust Vector Control System was required. CSD used the operationally proven Ball and Socket movable nozzle system for TVC. The motor utilizes a high length/diameter graphite epoxy case with full diameter nozzle attach opening, a high solids loaded HTPB propellant, and a high powered electro-mechanical actuation system for vectoring the nozzle. To assure flight tests at White Sands Missile Range (WSMR) stay within the confines of WSMR, it was necessary to use CSD's TVC system to put the THAAD missile through energy consuming maneuvers during the early portion of the flight. This is accomplished by introducing a corkscrew flight path immediately after launch. During the Lockheed Martin/ CSD proposal phase an all up prototype motor was manufactured and successfully static tested in order to demonstrate the system approach and mitigate program risk. US Army selected the Lockheed Martin/CSD team to develop and qualify the system. The motor went through a rigorous development, qualification, and flight test program. The propulsion/missile system has flown and successfully intercepted the target under various engagement scenarios.

23

THAAD Booster Motor Dramatic Demonstration of CSD’s Steering System Capability for Flight Energy Management Minuteman III In 1983 CSD was awarded a contract for the remanufacture of the 3rd stage of the Minuteman III (MM-III) ICBM. This program involved making new 52-in diameter fiberglass motor cases and loading them with typical 3rd stage propellant. The forward end of the grain had a 6-point star which gave access to six thrust termination (TT) ports for energy management. Much of the other inert hardware was taken from former 3rd stage motors which were removed from active locations. The program included remanufacture and test of the Safe and Arm devices, TVC valves and TT ordnance. Because of the success of this program, in the 1990s TRW and the Air Force decided to consolidate the number of propulsion suppliers involved in the 3-stage MM-III. Thiokol was given responsibility for the stage-1, CSD for stage-3. A Joint Venture (JV) was formed whereby Thiokol and CSD split the stage-2 responsibilities 50/50. All CSD-involved MM-III stage-2 and stage-3 motors flew with 100% success.

Minuteman III Stage-2 Minuteman III Minuteman III Stage-3 Launch A couple of missions took place using a modified version of the MM-III stage-3. This modified motor, called the Orbus 7S, was a spinner which used a MM-III loaded case, had a fixed nozzle and no thrust termination system. The Orbus 7 was used for space maneuvering type missions, such as providing the perigee burn for JC Sat and other payloads. The Orbus 7S performed flawlessly in its flights.

Hybrid and Ramjet Technology Hybrid Rockets CSD’s early experience with liquid propulsion systems (fuels, oxidizers, pumps, etc.) provided important insight into creating hybrid rocket propulsion systems which combined the attributes of

24

liquid oxidizers and solid fuels (which could be made from a broad range of materials including old tires, or even trash). From the 1960s into the 1990s, CSD was awarded dozens of contracts to demonstrate various approaches to hybrids and to advance the technology and viability of such systems. The hybrid programs which received the most press were: Sandpiper, High Altitude Supersonic Target (HAST) and Firebolt.

In the late 1960s, the Air Force flight tested some modified Beech AQM-37A target drones under the program called Sandpiper. The Sandpiper version removed the original hypergolic liquid propulsion system and replaced it with a CSD hybrid rocket which was inherently safer. The hybrid propulsion system was re-loadable (used solid fuel wafers) which was utilized by the program since the target drone was recoverable. After the successful Sandpiper tests, the Air Force formally started the XAQM-81A High Altitude Supersonic Target (HAST) program to develop a production target based on the AQM-37 (Sandpiper) configuration.

The HAST engine, built by CSD, was throttleable between 120-lb and 1200-lb of thrust. A ram air turbine, with an inlet below the center fuselage, pressurized the IRFNA (Inhibited Red Fuming Nitric Acid) oxidizer before it was delivered to the thrust chamber, and also provided electrical power for the missile. After air launch at about Mach 1.5 from an F-4 aircraft, the hybrid rocket could propel the XAQM-81A to speeds of more than Mach 4 at altitudes of 100,000 ft. The Firebolt could fly a pre-programmed course and/or respond to guidance commands from the ground. The parachute recovery system allowed either a soft landing or a mid-air retrieval.

Overall, CSD completed over 2,500 static tests of various hybrid propulsion systems.

A 1964 Hybrid Static Test at CSD’s Coyote Facility

CSD’s Hybrid Rocket provided Propulsion for USAF/Teledyne Firebolt Target Drone, including a CSD EMV for precise oxidizer flow control

Ramjet and Advanced Air-Breathing Propulsion Development In the early 1970s, ramjet development took place at both the Sunnyvale and Coyote Valley facilities. Sunnyvale was used for testing small and medium-scale ramjets with a test stand which

25

could deliver up to 35 pounds of heated air per second to simulate high Mach inlet air temperatures. Later, this stand was moved to the Coyote site. Two larger ramjet tests stands were put into operation at the Coyote facility. These test stands were initially used for liquid rocket testing and hybrid motor testing. A hot water supply system was installed between the two stations as a source of up to 75,000 pounds of steam for ejectors located at the two test stands. These stands and steam ejector system permitted testing at simulated flight Mach numbers of up to 6.0 and altitudes of up to 80,000 feet. These Coyote test stands were used for the testing of a variety of air-breathing propulsion systems: Solid Fuel Ramjets, Ducted Rockets, Liquid Fuel Ramjets, Air-Turbo Ramjets, Air-Turbo Rockets and Scramjet Endothermic-Fuel Systems. Several thousand direct-connect, freejet, semifreejet and boost-to-sustain transition tests were performed over 3 decades at CSD. In the course of these efforts, solid fuel ramjet grains were successfully demonstrated from -65°F to 165ºF while the liquid-fuel and ducted rocket engines were similarly demonstrated over a wide range of operating conditions. In addition to the air-breathing engine testing, the facility was used for thermal performance evaluations of materials and propulsion sub-systems and the testing of integral rocket ramjet boosters. In this regard, liquid fuel integral boosters were tested over the significant temperature range of -45ºF to 145ºF. Initially, the Sunnyvale efforts were focused upon solid fuel ramjet testing, while liquid fuel and ducted rocket ramjet testing was conducted at the Coyote site. CSD’s first major Liquid Fuel Ramjet (LFRJ) program was initiated in the late 60s in cooperation with the Naval Weapons Center at China Lake, CA. CSD participated in the development of the Advanced Low-Volume Ramjet (ALVRJ) liquid fuel ramjet. Development of this engine culminated in a successful flight test program in 1975-1976 with flights of up to Mach 2.9 and altitudes of 1,000 feet to 30,000 feet. In addition to the ALVRJ program, CSD was involved in the Modern Ramjet Engine (MRE), the Supersonic Test Missile (STM), the Advanced Strategic Air Launched Missile (ASALM), the Supersonic Low-Altitude Target (SLAT), the Advanced Large-Scale Swirl Engine, the Advanced Air-to-Air Missile (AAAM) and a variety of other liquid fuel ramjet programs.

A Portion of CSD’s Wing Mounted STM Ducted Rocket Propulsion Ramjet Test Facility Ramjet Missile Prototype Validation Test Vehicle

Numerous solid fuel ramjet (SFRJ) programs were conducted at CSD. These included both conventional solid booster/integral rocket ramjet missiles as well as engines and projectiles fired from guns. Major solid fuel ramjet efforts included the SFRJ Propulsion System Demonstration Program, the SFRJ Kinetic Round (SPARK), the SFRJ Advanced Indirect Fire System (AIFS), the RAMROD, the SFRJ Engine Demonstration (SFIRR), the SFRJ Boron Fuel Program, and a variety of gun-launched projectile programs (including the AIFS and RAMROD programs). These programs evaluated the potential of the SFRJ for a number of air-to-air and air-to-surface missile

26

applications. In the course of these programs, the validity of the solid fuel ramjet engine was demonstrated in engines ranging in size from 20-mm in diameter to 17-inches in diameter. While none of these ramjets or projectile systems was transitioned into weapon systems, there is ongoing interest in the potential of the ramjet to provide longer ranges and shorter times-to-target for a variety of applications.

Solid Propulsion Technology

Motor Components CSD was known both nationally and internationally for its innovative engineering. Besides design, engineering included extensive analyses to assure the viability of creative design concepts. Here are some examples of CSD’s exemplary engineering efforts:

1) Invented the idea of segmenting large solid rocket motor cases 2) Designed the clevis joint used to connect motor case segments 3) Designed and qualified the world’s first redundant driven electromechanical actuators to

vector IUS stage-1 & stage-2 nozzles 4) Created and tested the idea of a Bolt Extrusion Thrust Termination (BETT) using the

concept of extruding the nozzle attach bolts to control the aft movement of the nozzle during TT, thus eliminating dangerous reverse thrust spikes

5) CSD was the principal developer of the 3D/4D carbon-carbon integral throat and entrance (ITE). C-C ITEs significantly increased the robustness of nozzles, dramatically improving nozzle reliability

6) CSD was one of the first to use 2-D carbon-carbon for nozzle exit cones 7) CSD created (through IR&D funding) the “collapsible cup”, or “nested cones” extendible

exit cone (EEC) system which was later the first to be qualified for flight on the IUS 2nd stage

8) CSD creatively used sinusoidal beams to deploy the EEC nested cones on the IUS 2nd stage

9) CSD designed a nozzle environmental closure for the Shuttle BSMs which, upon opening, remains attached to the nozzle exit cone to eliminate ejecta while also serving as an aeroheat shield during Shuttle booster ascent

10) CSD did extensive in-house work to optimize filament winding of motor case designs, including the manufacture of one of the first motor chambers made out of Kevlar

11) CSD – often in cooperation with Snecma Propulsion Solide (SPS), Bordeaux, France – demonstrated numerous innovative applications of advanced composite materials. Notable were the all-composite motor demonstration, composite hot gas valves and chamber polar bosses, the integral throat and exit cone (ITEC) nozzle, and the IHPRPT-funded supersonic splitline flexseal nozzle

12) The design, test and eventual qualification of the world’s lowest torque movable nozzle system, the Techroll Seal, a constant-volume, fluid-filled (silicone oil) bearing which incorporates rolling convolutes to virtually eliminate resistance in the system during omniaxial vectoring

13) The pressure curing of propellant for motors using filament wound motor cases in order to provide near stress free propellant grains during storage.

Additionally, CSD was highly respected for its hardware forensics capabilities. This involved the evaluation of both successful and failed hardware. The analysis of successful hardware assured that no incipient conditions might exist which would cause failure when the motor operated at conditions other than those tested to-date. The analysis of failed hardware was critical to getting a program back on its desired path of success. This highly successful CSD capability was sometimes utilized by outside organizations when they experienced a failure.

27

An excellent example of the transfer of solid rocket technology to liquid engines is the case of the EEC nested cones used on the IUS program which resulted in the Orbus-6E of IUS delivering the highest specific impulse (304.3-sec) for a flight qualified solid rocket. CSD engineers worked closely with the engineers at Pratt & Whitney’s Space Propulsion Division in West Palm Beach, Florida to understand the needs and challenges of their new version of their significant line of RL10 cryogenic engines. The translatable carbon-carbon cone approach was subsequently incorporated into the nozzle design of the RL10 B-2 upper stage of P&W-WPB. The result is that the RL10 B-2 delivers the highest specific impulse of any flight qualified liquid propulsion system in the world.

IUS’s Orbus 6E SRM pioneered C/C extendible nozzle technology, the design concept was transferred to P&W-WPB’s RL10 B-2 which is the

upper stage of the Delta IV Launch Vehicle

Propellant Development and Analysis CSD’s Research Department worked hand-in-glove with CSD’s Engineering Department in order to make sure the needs of engineering’s designs were met and to make sure the best possible understanding of the operation of the engineering designs was obtained. CSD used the “team” approach to solve problems and challenges. An excellent example was the Operational Propellant Committee (OPC) which included representation from Engineering, Quality Control and Operations (propellant processing). The OPC had significant influence over all propellant operations.

---Propellant Formulation Development: CSD was instrumental in developing PBAN formulations for various booster and space applications. This evolved into the development of HTPB formulations for many space and missile applications to take advantage of the improved physical properties and performance attributes of this propellant system. Additionally, CSD pioneered the use of HMX as a performance improvement additive to HTPB-based systems for use in upper stages. CSD made many contributions to the solid rocket industry in the development of propellant processing aids and propellant stabilizers. An example was PROTECH® which appreciably improved aging characteristics. Similar work was carried out by CSD for hybrid rocket fuels and oxidizers and fuels and oxidizers for solid fueled ramjet applications.

28

CSD was deeply involved in the microscopic assessment of solid propellant during its combustion process. This resulted in an important insight as to what occurred during combustion: measured aluminum agglomeration size distribution, flame propagation into narrow slits and the behavior of propellant combustion under acceleration, such as that resulting from motors which spin. Knowing how a motor’s propellant burns under acceleration loads was critical to the development of propellants for such CSD motors as FW-4 (up to 200 rpm spin rate) and the Orbus 21S for Intelsat VI. The Orbus 21S contained a propellant load of 21,260-lb. When spun at 30 rpm during a static test, the Orbus 21S retained only 30-lb of slag, or 0.14% of the original propellant load—a very low value.

Propellant combustion studies provided important data concerning ignition and flame propagation insight. Laboratory T-Burner test results, and rotating valve analyses provided data about the potential susceptibility of the system to experience combustion instability.

Microscopic photo of propellant sample burning: HTPB propellant containing HMX. It shows burning aluminum particles, some as large as 600 to 720 microns

---Flow Analysis: A computer generated analysis for flow within the combustion chamber and within the nozzle is critically important in order to more fully understand the characteristics of the motor before it is ever built or tested. These analyses can be used to establish pressure distributions on the propellant grain, pressure drop in the motor and help optimize the nozzle contour design to possibly minimize length or possible particle impingement on the nozzle wall. This type of analysis allows parameters to be varied in order to ultimately end up with a motor near its optimum design point.

Important flow analysis tools were computational fluid dynamics (CFD) computer programs. These programs could look in detail at flow conditions inside a motor. They could predict whether or not vortices would form. Using these tools, one could modify a propellant grain configuration to minimize its propensity to create vortices which might create a combustion instability problem.

---Computer Performance Analysis: CSD was intimately involved in the creation of a standard propulsion performance program (SPP) for the Air Force Rocket Propulsion Lab (AFRPL). This program incorporated CSD’s 3-dimensional grain design and internal ballistics programs in conjunction with a 2-dimensional, 2-phase performance loss program to provide an overall motor performance analysis program. Additionally, this program was “married” to the motor case structural analysis program so that the effect of case growth during pressurization could be translated to propellant grain dimensional changes. AFRPL made this program available to other organizations with a need-to-know within industry and academia. This program became and still is an industry standard.

---Non-Destructive Testing (NDE): A major product of the Titan Recovery Program was a comprehensive SRM NDE program developed in partnership with experts at United Technologies Research Center (UTRC). This included state-of-the-art internal imaging systems using Ultrasonic, Laser, Infrared and X-Ray technology. Facilities were built at both Titan Launch Bases as well as CSD’s Coyote Facility. As part of the Navy’s Trident II (D5) program, an

29

advanced in-ground Computed Tomography (CT) facility was built at CSD’s Coyote Facility. This facility was also used for CSD’s MM-III and BSM motors.

Electric Propulsion

In July 2000 CSD acquired Space Power, Inc. (SPI), located in Sunnyvale, CA. SPI was involved in electric propulsion systems, including Hall Effect Thrusters (HETs), power processing units (PPUs), propellant management systems (PMSs) and propellant tanks. During the ensuing years, CSD worked closely with the University of Michigan and NASA-Glenn to test a series of HETs, demonstrating their impressive specific impulse (Isp) capabilities as well as durability. CSD worked closely with several satellite companies to make sure CSD’s efforts went along a path to complement the needs of potential customers. The eventual shutting down of the CSD facility terminated this effort before it could reach fruition.

T-220 HET 1000-hour test at NASA Glenn: demonstrated a 2,500-sec Isp

Outstanding Demonstrated Product Dependability

During its history, CSD produced a wide variety of SRM propulsion systems. Over 4,000 of these SRMs were involved in operational flights. Only three of these SRMs had failures which resulted in loss of the mission, and one SRM had an anomaly in which the payload eventually reached its intended orbit. These results provide a demonstrated dependability of less than one failure incurred for every 1000 SRMs launched.

Similarly, CSD was noted and respected for its extensive capability to design, manufacture and test electromechanical controls, valves and avionics devices, especially to support the very demanding steering needs of various flight vehicles. This area of CSD had a perfect record of dependability for over 5,000 avionics systems used on operational flights.

A Quick Look at Some of CSD’s Leaders

With the retirement of Gen. Putt, Mr. Barnet Adelman became the President of CSD (then United Technology Center) in 1962 and led the company through its critical early Titan efforts, resulting in the 5-segment Titan boosters which demonstrated a 100% success flight record. He was also responsible for the development of the Titan 34D and Titan IVA SRMs. Mr. Adelman guided the expansion of CSD’s business base beyond Titan to include such programs as FW-4, Algol III, FW-5, Shuttle’s Booster Separation Motors and the Inertial Upper Stage (IUS).

30

Mr. Eugene Roberts was instrumental in making CSD’s Coyote Development Center into a world-class propellant processing, hardware manufacturing and motor test facility. His efforts were key to the success of moving the Titan IIIC SRM program forward on such an impressive schedule. Later, Mr. Roberts became the VP of all Programs at CSD, demonstrating his great ability to work well with CSD’s customers as well as CSD’s employees.

In 1980, Mr. Adolph “Al” Medica was brought in from UTC’s Norden Systems to succeed Mr. Adelman who was elevated to a UTC Corporate Vice President position. Under Mr. Medica’s leadership, CSD expanded into the Tomahawk Mk-111 Program, the Intelsat VI use of CSD’s Orbus 21, and the start of the Minuteman III Remanufacture Program. Mr. Medica was also responsible for eventually consolidating all San Jose area facilities into one location: the Research, Engineering, Development and Test Facility in Coyote Valley.

From CSD’s earliest days, Dr. Bernard “Ross” Felix was an important leader of the Engineering evolution of CSD. With a superb knowledge of propulsion systems, Dr. Felix was the key technical figure guiding many of CSD’s programs. After a time as Engineering Department Manager, Dr. Felix took over the new position of VP-Engineering. Dr. Felix became an Engineering icon within the solid rocket industry, gaining the highest respect of his CSD associates, his customers and his competitors.

Mr. Douglas North was one of the very early employees of CSD, working his way up from design engineer, to Engineering Manager, to VP-Engineering, to CSD General Manager, to President of the newly formed Pratt & Whitney Space Propulsion (a Division of Pratt & Whitney). P&W-SP included the solid rocket site of San Jose, CA and the liquid engine site of West Palm Beach, FL. Mr. North’s technical leadership was critical to the success of virtually every solid rocket program at CSD, as well as the Titan Recovery Program after an SRM-caused flight failure. His management leadership was critical to establishing a supportive/synergistic relationship between the San Jose and West Palm Beach organizations.

Mr. Charles Sinclair also started at CSD during its very earliest days. Working as an engineer in both the test and design areas, Mr. Sinclair developed an in-depth understanding of rocket propulsion systems. He worked his way up through various positions to Engineering Manager and later became CSD’s VP & General Manager when Mr. North became President. Besides his engineering skills, Mr. Sinclair was noted for his negotiating prowess and was instrumental in negotiating the consolidation contract for the MM-III program. Mr. Sinclair was also a critical leader in establishing the teaming arrangement with Lockheed-Martin which eventually won CSD the booster program for the THAAD missile.

31

The Closing of a Propulsion Icon

In the late 1990s and early 2000s, a number of solid rocket programs were being consolidated in order to minimize the number of propulsion contractors involved. Included in these programs were Trident II (D5), Minuteman III and others. During this period of time, several attempts were made for CSD to merge with other solid rocket companies in order to consolidate facilities and personnel, but to no avail. CSD also had many employees with extensive years of service with CSD (some over 40-yrs), so that retirement of these personnel was imminent. Bringing in younger personnel was becoming more and more challenging because of the very high cost of living which had evolved in the San Francisco Bay Area.

When originally constructed, the Coyote Facility was located in the remote canyon lands south of San Jose. By 2003, this facility was being encroached upon by housing developments, which placed additional restrictions upon its operation.

By the early 2000s, CSD had long established itself as one of the preeminent solid rocket companies of the world. From 1958 to 2003 CSD had been directly involved in most of the key space and missile programs of the United States. At its 5,400 acre processing facility, CSD had processed over 300 million pounds of propellant and conducted thousands of static motor firings in state-of-the-art facilities. All propellant was processed remotely, in order to maximize safety.

On 7 August 2003 propellant was being processed in one of CSD’s 600-gallon mixers. Something within the mixer created an energy source which ignited the propellant, destroying the mix station. Immediately, an investigation team was formed to determine the cause of the problem and what actions were needed to get CSD back on the path of processing propellant. As part of the resulting action, CSD decided to go through the other mixing stations, clean all aspects of the equipment and make any necessary repairs or updates. During this process, work on some equipment at another mix station caused an explosion resulting in a fatality—the first ever at a CSD mixing facility. After many very high level discussions, the corporation decided to permanently terminate motor processing operations at the Coyote Valley Facility.

At the time of these problems, CSD was actively involved in producing rockets for many critical government space and missile programs. Following discussions with our customers and their associated government agencies, it was decided that CSD would transition all of its programs to other appropriate propulsion companies, as directed by the government. A large group of highly devoted CSD employees elected to stay at CSD for about 18 months (and in several cases about 2 years) to make sure that this transition process went very smoothly, and that the products produced by the transition supplier provided performance identical to the products made by CSD.

As part of the closing process, CSD maintained a special staff at the facility to clean any potential contamination and to dismantle all buildings. The final use of this site has not been announced.

Acknowledgments

CSD was a wonderful company for which to work. Over the years, many thousands of employees made this company extra special, creating a close knit family whose mutual support gave our workplace a wonderful environment. CSD had a remarkable collection of creative minds that were key to the design innovations for which CSD became known. CSD’s cadre of people in all functional areas contributed in their own creative ways to ultimately make our systems outstandingly dependable. It is these people, whose dedication and hard work deserves a hearty thanks and congratulations. This history paper is dedicated as a tribute to these very special associates and to provide them wonderful memories of their historic achievements.

32

I would like to thank Mr. Barnet Adelman and Dr. David Altman for the insight they provided me about the very early history of CSD.

With respect to the actual preparation of this document, I would like to thank the following people for the very important inputs they provided me regarding the various programs. They are (in alphabetical order): Ron Adib, Stanley Backlund, Darryl Bramlette, Robert Bresniker, Russell Ellis, Ralph Hammond, Frank Inman, Dr. Gordon Jensen, Dr. Frank Lavacot, John Lee, George Morefield, Douglas North, Carl Pignoli, Harvey Sakai, Paul Willoughby and Frank Verlot. Of these, I would especially like to thank Mr. Backlund, Mr. Bresniker and Mr. Ellis for their many reviews of my document and their invaluable inputs which were critical to the accuracy of this document.

Footnote

CSD had a successful history as an important supplier of solid rocket propulsion systems for many critical US space and missile programs. This success benefited from working with other organizations within the Air Force, Navy, Army and NASA. These organizations were, in turn, supported by important consulting organizations and a long list of prime contractors. Success would have been impossible without support CSD received from the many talented and dedicated subcontractors and suppliers of critical components and hardware. Additionally, CSD enjoyed the unique position of help, when needed, from its sister divisions within United Technologies such as Pratt & Whitney, Corporate Systems Center (CSC), Sikorsky, Hamilton Standard (Sundstrand) and United Technologies Research Center (UTRC).

Disclaimer This paper was written by the author with the support of those persons mentioned in the above “Acknowledgements” section. All of us are retired employees of CSD and, therefore, do not officially represent CSD. The purpose of this paper was to record the history of CSD and its pioneering propulsion accomplishments while also paying tribute to some of its key leaders who were instrumental in the development of the company. This paper was not solicited, supported or sanctioned by CSD, Pratt & Whitney or United Technologies Corporation. Additionally, I disclaim any agency, contract or employee relationship with UTC or its Pratt & Whitney business division during the preparation of this paper.

33