I A STATUS REPORT LOCKHEED

I for the 7th ANNUAL AIM-UTAH STATE UNIVERSITY I CONFERENCE ON SMALL SATEWTES D.E. Davis, J.W. AngelP, A.J. MaeLaren2 I Lockheed Missiles & Spaee Co.,lne Abstract: This paper discusses a new collocated product development team family of small and medium space launch with the mandate to use our engineering, vehicles being developed by Lockheed manufacturing, and flight test I Missiles & Space Company, Inc. The experience, but - and an important but development program will culminate in - to be innovative in our approaches. a demonstration launch in November For example we elected to use, as much I 1994. The paper gives a brief as possible, existing hardware; to use background and gives the program status aluminum structure rather than as of the date of this paper. composites - although Lockheed builds Supporting graphics are included. much composite structure; and to I simplify the way we do business. Background: We have streamlined our Lockheed Missiles & Space Co, paperwork system. The product I Inc. has been studying small space development team has informal reviews launch vehicles since 1987. The - informal in the sense that their are no original approach then was to use dry-runs, and hand drawn graphics are I surplus or excess ballistic missile acceptable. The technical caliber of assets, primarily the 1st and 2nd stages the presentations is as professional as if from the Poseidon C-3 missile as the it were our traditional government basis, replacing the weapon system customer instead of just ourselves. I guidance and control with state of the art Rather than tell our prospective technology, and adding a 3rd stage with subcontractors how they should design a 48 solid rocket motor. Although their hardware, we have either accepted I the approach was notionally attractive their specification or we have discussed and the calculated reliability was over our requirements and allowed them to 90%, no users were found. suggest the solutions and write the Motorola's announcement of the specification. We have held an oral I Iridium™ program in 1990 competition where the proposal was the dramatically altered our approach. viewgraphs and a one page letter on Motorola's imperatives were very clear: price. This has not been without I lowest possible cost, on demand launCh, some internal trauma and there have and outstanding reliability. When we been replacements of personnel who worked through the alternatives, we could not cope with the new world order. I concluded that reliability had to be as On January 8, 1993 Lockheed close to 100% as possible. A 90% to Corporation approved a company funded 95% reliability was unacceptable; we development program and demon­ calculated that for program such as stration launch in November 1994. I Iridium™ we had to self-insure, because We have submitted the required range the costs of insurance at 15 to 18 documentation to both the Eastern Range percent per launch would be (Cape Canaveral) and the Western Range I unacceptable from a price strategy point (Vandenberg Air Force Base). The Air of view. Thus we quickly discarded the Force has approved our use of Space ballistic missile assets approach, and Launch Complex (SLC-6) at Vandenberg I reexamined approaches based on AFB. We are working with the Eastern 's 120™ which was then Range and the Florida Space Port I in development. We formed a Authority at Cape Canaveral. We have I 1. P.E., Member ASME 2. Senior Member AIAA I 2 started discussions with the Office of at the launch site to final countdown and I Commercial Space Transportation launch is a short as two weeks. (OCST) of the Department of The first of these vehicles is Transportation (DoT) with respect to scheduled to fly from Vandenberg in I the relevant licenses required by the November 1994. General operations Commercial Space Launch Act (CSLA). will begin in 1995. We completed our Preliminary Design Review (PDR) process in July, Solid Rocket Motors: I and have started fabrication of the first The LL V1 first stage is Thiokol hardware designated for testing. We Corporation's Castor 120™ Solid Rocket are on a clear track to meet the Motor. It is a 120,000 Ibm class motor I November 1994 launch date mandated that employs a graphite epoxy composite by our CEO, Dan Tellep. case, Class 1.3 HTPB propellant, pyrogen igniter, and a vectorable nozzle Launch Vehicle Description: driven by a cold gas blow down thrust I This section describes the design vector control system. This motor is and system performance of this new used as both the first stage and the family of launch vehicles designated the second stage on the LLV2 and LLV3. As a I Lockheed Launch Vehicles. second stage motor, a larger expansion The smallest vehicle (LLV1) will ratio nozzle is used in order to improve place about 2200 pounds in to a low performance at higher altitudes. The I Earth orbit (LEO) of 100 nautical miles propellant grain can be tailored to at 28°. The next increment, (LLV2) change the burn characteristics and will place 4000 pounds into LEO, and thrust history. This provides the the LLV3 will place up to 8000 pounds ability to taper the thrust of the second I into LEO. Launch operations will be stage near burn out, which lowers the conducted from Vandenberg Air Force vehicle acceleration and provide~ a Base in California for Polar and smoother ride for the payload. I Synchronous orbits and from Cape The upper stage of the LLV family Canaveral. A mobile launch system is the Orbus 21 D solid rocket motor checkout van will be provided which manufactured by the Chemical Systems will allow launch operations from any Division of United Technologies. Inc. I suitable range. The Orbus 21 is a proven motor which Launch Vehicle assembly and has been used in the checkout at the launch site is based on a on 17 flights. For use on the LLV it will I simple stack and shoot approach allowed use a carbon phenolic nozzle and an by the inherent simplicity of solid electromechanical actuator. rocket motors. Only the Castor IVATM solid rocket motors I System, which will be fueled prior to may be strapped on to the LL V family of assembly on the pad, uses liquid vehicles to increase payload to orbit propellant. The payload () capability. will be mated to the payload adapter and I release mechanism in a clean room. Attitude Control System After payload checkout, the Fairing will The Orbital Assist Module(OAM) be installed and the resulting is located above the Orbus 21 D. It I Encapsulated Assembly will be moved contains the LLV Flight Electronics, and placed on top of the waiting solid Batteries, Telemetry, Inertial rocket motors on the pad. Final Measurement Unit, and the Attitude preparation and check out on the pad Control System (ACS). The ACS is a I includes installation of batteries, liquid monopropellant hydrazine destruct arming devices and final propulsion system which has 10 rocket electrical checks. The time from engine assemblies for pitch, roll, yaw I completion of payload receipt inspection control and velocity addition to correct for any errors induced during solid I I I 3

I rocket motor boost flight. The ACS may of altitude and 0.020 inclination of the be configured with 2, 4 or 6 hydrazine desired orbit. tanks so that the propellant load may be I tailored to the specific mission. Fuel Payload Accommodations load with the modular tanks is 260, Of specific interest to the 520, and 780 pounds. The ACS is potential user of a new launch vehicle manufactured by the Rocket Research are the payload accommodations I Company which provides four 50 Ibf features. These accommodations consist axial thrusters for velocity addition and of the fairings, payload adapters and six 25 Ibf thrusters for pitch, yaw and release mechanisms, on pad air­ I roll control. The hydrazine tanks which conditioning, T -0 umbilical and support are spun aluminum with a graphite equipment interfaces and electrical composite overwrap, are pressurized interfaces provid i ng te Ie metry with gaseous nitrogen to 460 psig and monitoring capability before and during I contains a elastomeric bladder for flight. The LLV offers some unique positive expulsion of the fuel. The ACS features in this area. system may use up to 98% of the fuel on The LLV family has been designed I board at an average of with the underlying theme of reducing 220 seconds. The total impulse system cost. This is especially reflected available is 57,200 Ibf sec. All in the selection of available fairings I components of the ACS have a proven range from a 92 inch diameter fairing flight history which minimizes intended primarily for the LLV1 to the development cost and program risk. 141 inch fairing for the larger vehicles. The fairings are of a two piece clam I Guidance, Navigation & Control shell design. Zip tube, a patented clean The heart of the LLV Guidance, system, cuts the base and two halves Navigation and Control (GN&C) system allowing springs to provide the I is the Litton LN-100L Inertial separation forces. The increased volume Navigation Unit (INU). This is a strap of these fairings, which are large for down system which employs a this class of launch vehicle, simplifies nondithered Litton Zero-lock™ Laser the packaging requirements for I Gyro(ZLGTM) and A-4 accelerometer deployables such as solar arrays and technologies together with a antennas. There is, however, a sophisticated 24-state Kalman filter. performance penalty due to the I This is a low cost, advanced technology increased weight and drag of the larger INU with flight demonstrated fairings, therefore the small fairing is performance of 0.4 to 0.6 nmi/hr designed for use with all three vehicles, (CEP) which validates simulations and while the 116 inch diameter fairing and I exceeds SNU 84-1 type medium the 141 inch diameter fairing can be accuracy requirements. This unit is used on either the LL V2 or LLV3. Should derived from a commercial unit and has a payload nearly fit within an envelope, I been designed for Military aircraft and but violate it at a few isolated locations, launch vehicle requirements. Less than specific accommodations are potentially 500 cubic inches in volume, it weighs available. The fairings are equipped I 18.5 pounds and uses 27 watts. with access doors and RF windows which Predicted mean time between failures is are located per the satellite greater than 5,500 hours during boost requirements. phase and 26,000 hours during space The approach to the payload I flight coast phase. No scheduled interface adapter design maintains the calibration is required. The sensitivity flexibility required to meet varied needs of this unit allows the LLV to place a of payloads. It features a fixed conical I typical payload within one nautical mile adapter section reducing the 92 inch diameter of the equipment section to a I I I 4 standard 66 inch bolt circle. This There are commands which can be I interface is compatible with existing generated by the flight computer in payload adapters and separation accordance with satellite speCified mechanisms which several satellite criteria. These can be used to initialize I manufacturers have developed for use satellite inertial reference systems. with other systems. For smaller open shutters, or similar spacecraft . the cone can be extended to functions requiring actuation prior to reduce the diameter to be compatible separation. I with the 38.8 inch marmon clamp system frequently used. Alternatively. Flight Environments we will offer a range of adapters based The environments experienced I on the requirements of the early flights. by the payload during flight on a LLV are For vehicles using the larger fairings. similar to those of other launch the equipment section diameter can be vehicles. Because the LLV design I extended upward to a 92 inch diameter utilizes existing solid rocket motors, the interface providing a more efficient load axial static load is larger than for some path. The mass properties book keeping liquids. It varies depending upon ascent procedure assigns the weight of the trajectory and payload mass but does not I adapter and separation mechanism up 10 exceed 8g. The lateral load maximum of the payload interface to the launch 2.5g is below those of winged launch vehicle for purposes of performance vehicles and is comparable to liquids. I calculations. Adjustments will be made The dynamic loads associated with motor should the satellite choose to provide the ignition and stage separation are below adaptor/separation mechanism. these limits and are not additive to the Umbilical connectors, which are static load. The stiffness required of the I pulled at separation. provide the satellite and the adaptor interface is 15 electrical interface to the LLV avionics Hz for the LLV1 to avoid coupling with and a pass through to the T -0 umbilical. the structural modes of the launch I For spacecraft T-O umbilical functions. vehicle. The requirement for LLV2 and there are 74 shielded copper circuits of LLV3 will be lower due to the increased several wire gauges. These circuits are vehicle length but have not yet been I carried through the equipment section calculated. and the flyaway umbilical to a utility Acoustic environment for the room located beneath the launch pad. LLV1 was developed using measured data From there they may be routed either to obtained during test firings of the Castor I the launch control van via a two way 120™. our database from firings using fiber optic data system or 10 local ducted pads similar to those planned for equipment per the satellites LLV. flight data from similar fairings I requirement. Routed along with these and the VAPEPS computer code. The circuits are two fiber optic cables which relatively low level derives from the can be used by the satellite to provide a small size of the vehicle and the use of a high data rate path with the satellite. ducted pad. The acoustic environments I These can be used for memory for LLV2 and 3 have not yet been verifications and other operations which estimated but are anticipated to be are frequently limited by the bandwidth similar or lower due to the increased I of the launch pad data link. They include height of the payload at launch and the analog monitors for battery voltages. lower maximum dynamic pressures pressures. etc. discrete on/off during flight. I indications. and signal conditioned The pyro shock environment is continuity lops which may be routed developed at separation of the nose through the satellite to provide reset or fairing. ignition and separation of the safe loop indicators. These are in Orbus 210 and payload separation. It is I addition to the separation indicators based on measured test data for the which are dedicated to that purpose. I I I 5

I respective sources scaled to reflect the The LEO payload capability to coupling path. various inclinations for the LLV1 is shown in a following figure. The effect I Stack and Shoot Approach of adding the second Castor 120™ to One of the key advantages of the make an LL V2 is also shown. Since the minimal complexity of the LLV payload interfaces and much of the manifests itself at the launch base. The vehicle are common for all I modular design produces a minimum configurations, this provides a satellite number of systems which must be developer with a new option should his integrated and verified. The solid rocket payload increase in weight beyond the I motors and interstages are shipped capabilities of the intended launch directly to the launch pad for assembly. vehicle. For the minimal incremental The Orbit Adjust Module is checked-out cost of adding the second stage , at the factory and delivered directly to approximately 2500 additional pounds I a propellant loading facility where the of capability is available. This often is ACS hydrazine propellant is loaded. less than the cost of a weight reduction Using methodologies developed to process program or compromises in satellite I ballistic missiles (over 1100 capability. The concept of incremental launches). the assembly process of an performance increases continues with LLV1 is planned for a span of 14 days, the LLV3. Castor IVAs are added in I the LLV2 and LLV3 taking slightly quantities of 2, 3, 4, and 6 to provide longer. performance increases of 1400, 660, Requirements for interactions 600, and 1010 lb. respectively. during payload and launch vehicle I processing have been minimized. The Summary payload processing span can be as long The new family of launch or as short as necessary to meet the vehicles described in this paper is being I payload requirements. When ready, it is developed by a team led by Lockheed mated to the payload adaptor and fairing Missiles and Space Company. The design are installed in the payload processing maximizes the use of existing hardware facility. The encapsulated payload is to lower developments cost and shorten I then delivered to the pad for mating with the time between concept and production. the launch vehicle just 3 days before The first demonstration flight is launch. Filtered air-conditioning scheduled for 1 November 1994 form I maintains both the temperature and Vandenberg Air Force Base in California. cleanliness of the payload on the pad. The design is based on a modular concept where solid rocket motors are assembled Performance in various combinations to cover a range I Two profiles for injecting a of payload requirements from 1000 Ibm satellite payload into orbit are available. to 8000 Ibm to LEO. This approach Direct injection is applicable to low allows the user to select a propulsion I altitude circular or parking orbits and system that just fits his needs. It also for elliptical orbits. If a higher altitude allows the user to move up to a slightly circular orbit is desired, the indirect larger launch vehicle if a change in I injection profile provides greater mass Mission requirements causes payload to orbit capability. With this technique, weight to grow. This can be the satellite is directly injected near accomplished without facing completely perigee of an elliptical orbit with apogee new design interfaces and environments I at the target orbit altitude. When apogee caused by moving up to a new class of is reached, one-half orbit later, the launch vehicles. OAM thrusters are used to raise the Solid rocket motors also provide I perigee to provide a circular orbit at the a great deal of operational flexibility. targeted altitude. They are ready to launch when the I payload is ready. The stack and shoot I I 6 approach minimizes the time spent at I the launch site, lowering cost for the launch vehicle as well as the cost of the payload final assembly and check out I team. The launch vehicle is ready and waiting on the payload rather than the other way around. I Bibliography:

Solid Rocket Motor Space l a u n c h I Vehicles. A.J. Maclaren & H.D. Trudeau, Acta Astronautica Vol 30, pp. 165-172, 1993 Pergamon Press Ltd I Lockheed Thinks Small Donald F Robertson, Space, June 1993, pp.24- I 26

Ball Drops Pegasus Plan. Selects Lockheed Launcher, James R. Asker, I Aviation Week, Jul 19, 1993 I I I I I I I I I I I ------____ 1liii' _

mso LLV1 PERFORMANCE ~Lockheed

Payload .... 100 kg 5 6 7 8 9 10

_ i=99° 700 i = 90° " ""'. "\ .", _ i=57° 1200 600 """""'. '~", _ i = 28.5° • ...-4 ""'.". \'" § 500 ,,',,- \.. ", ", 900 >- ;:t....,..--- l " \" s:: OJ 400 0.. ""d ".. ".. '" ~ (l) .2 .. 600 l ...... '.0 " " 300 " ".. S -< ", ".. , ~ 200 " .. --::"';'- ... ------~~ 300 ----_.---...... :;:-.. ,.... " '~---- ".. --- Circular direct ------:.'"':..::'" _::", ---"__ -- 100 Elliptical direct ---Circular via 180° transfer 0 10 12 14 16 18 20 22 Payload,... 100 lb

M9076F3011 1 mSD LLV2 PERFORMANCE ~lockheed

Payload .... 100 kg 12.5 15 17.5 20 22.5

_ i=99° 700 i = 900 _ i=57° 1200 600 _ i = 28.5 0

...... § 500 900 >- -0'. I 2" OJ 400 , p.. "'0 ... " (I) .2 600 ~ :Jj ,,'.. ~ 300 ", S < " .. 200 ""'''.. " 300 ------.. -- -.. -.. ... ------" ".... --- Circular direct ------100 Elliptical direct ---Circular via 1800 transfer 0 30 35 40 45 50 Payload,... 100 lb ------M9076F3012-- 1 ------~

~Lockheed mSD LLV3~)PERFORMANCE

Payload -- 100 kg 25 27.5 30 32.5 35 37.5 40

_ i=99° 700 i = 900 _ i=57° 1200 600 _ i = 28.50

.- 500 900 ~,..,. § ,..,. s::-. (J) 400 p.. "0 (1) .E I, '.0 600 l 300 -« '. '-, '-. S ...... --- ...... ~""'""ii". ~\ ...... '- ""- " 200 -... ,-. " ...... '""" ...... '...... ' .... , .. -...-- ...... ~ -- -...... >~._..." "'-'''' ~...... ----...... ~ , 300 --- Circular direct ...... : •.:~ -...... :;...... 100 --Elliptical direct --- Circular via 1800 transfer 0 55 60 65 70 75 80 85 Payload,.., 100 lb

M9076F3016 1 mSD INDIRECT INJECTION PROFILE ~Lockheed

Circularization burn

II V trim burn

Orbus burn

First (and second) stage burn

------M9076F3002 1 ------~-

mSD DIRECT INJECTION PROFILE ~Lockheed

dV trim burn

..-.1--- Orbus burn

Coast

First (and second) stage burn

M9076F3009 1 PAYLOAD FAIRING mSD DYNAMIC ENVELOPES ~lockheed

92-in. FAIRING 116-in. FAIRING 141-in. FAIRING

298 In. 269 in.

78in.-.1 · ~ 103 In. -.1 · /---- 127 In. -.1

------M9092F3017 1 ------.-

mSD FIXED ADAPTOR SECTION

... 66.00 ------_tll_...

14.00

1...... 1------92.54~------tll-...

7075 T7351 ROLLED SHEET

DETAIL A DETAIL B

M9110F3001 1 mSD LLV ELECTRICAL INTERFACE ANALOG TELEMETRY MONITORS CONTINUITY LOOPS (10 PROVIDED) (5 PROVIDED) LAUNCH VEHICLE f SATELLITE LAUNCH VEHICLE I SATELLITE I

OT05V ®~

SEPARA TION INDICA TOR DISCRETE TELEMETRY MONITORS (5 PROVIDED) (10 PROVIDED) I I

1 Mn <2V=O >2V=1 COMMANDS (8 PROVIDED)

1 Mn

COMMAND ISSUED PER CUSTOMER· SPECIFIED CRITERIA M9092F3015 3 ------. - LL V1 LAUNCH ACOUSTIC mSD ENVIRONMENT PRELIMINARY

OVERALL SOUND PRESSURE: 136.2 dB

; 150 ! ! I I INTERNAL TO FAIRING I I- 145 - I • NO ACOUSTIC BLANKET I • ! I- • 1/500 LEVEL I 140 - I I • DUCTED EXHAUST ! -r.o -_. -- '0 135 - . . I . -. "t- - ! W j ...... ---t-_- I-- .- > 130 L[-- L- --r-_~L_",,_, 1 -_ ct ! I I- -1 0 125 - ...... _. __ ...... _...... •...... 0 ------... -- -- Itl ~k ,... ----..,. ······l en 120 ~~ L -T"" I I ..J ~Vl a. 115 I"" U) I~~ 110 -. -I--' ... -1-

._. .... I -·1·- I- _. --- ~--"'--' --- -- 105 ~r 100 -\- L 10 100 1000 10000 FREQUENCY (Hz)

M9092F3004 3 LLV1 RANDOM VIBRATION mSD ENVIRONMENT PRELIMINARY OVERALL LEVELS BOOST PHASE 7.89rms LAUNCH PHASE 3.1 9rms

• AT PAYLOAD INTERFACE • 1/500 LEVEL

0.01

--J-. ~. --:-:: ...-:-. :-: '-'1'" ... N -J: ---····lAUNCH PHASE··· '"-C) -0 en a. 0.001 ,--,+-1 1--+-1----1-- - - ..

1------1----1- -+·-+-I-+-t----t--

10 100 1000 10000 FREQUENCY (Hz) ------M9092F3005 3 ------. -

mSD LL V1 SHOCK SPECTRA PRELIMINARY

10000~~~~~~~~~~~~.. ~ .. -.=-.~-~-.-~-~-~~+~ .... ~-.=-.=--..=-= .. -.. ~... -=.--~---.~ .. ~~--~-+~.- .•.~ ..... • AT PAYLOAD INTERFACE .:~ .-::::::::... ~:::::::.=. :.::::t::::.. -::.':.'.==.:.::::.~ .. :::==: r::=.: :.=. r::::.~.~.~.. ':::f.: :.~ • LLV1 • 1/500 LEVEL 'ii) , i-:--...... •.... --.... ---... .p - -- .--.. -.-. -

~ 1000~==~==+=+=t+~~==~==+=+=~~~~~~~~~~~ i' :::::::::::::::::::::::::: ::::::::::::::: :::::::::: ::::::::: .::::::: :::::(:: :::: ::. ::: ... :::::::::::::::::::::::: :::::::::::::::: :::::::::::;;;;;;-. ~rF: i t:.:::::::::::=::::: ::::::::: ::::::::::: ::::: ::::::: :::::: :=- ::: ::. ~===...... - ...... ~ ==~~~~ =~ ~.~ =:== ...... ~ ...... ·t .. ·...... ~'.

~ . ------[!.- +---V- -71 --. ... --- .. - '------.. J_ .- ... ~ 100 -=::::::= = co::::: ::::::::::::1::,.:: ::: :=::::;i:=:::: ~ co:: .::: :::: ~ ~~~~OF~R~~~~~~~~~TION :: o • IGNITION EVENT <{ ~---t--+-----+--t-- =: - ~~ ....~--1 -/~-- -==+=t:=t=~=...... - .... - .....- ...... '" !=

10 100 1000 10000 FREQUENCY (Hz)

M9092F3006 2 mSD STACK-AND-SHOOT APPROACH L-DAY -15 -14 -13 -12 -11 -10 ·9 -8 -7 -6 -5 -4 ·3 ·2 -1 0

PAYLOAD PROCESSING FACILITY

PAD

------M9092F3016 5