The Space Congress® Proceedings 1983 (20th) Space: The Next Twenty Years

Apr 1st, 8:00 AM

Processing and Deploying the McDonnell Douglas Payload Assist Module (PAM)

C. E. Bryan Electronic Systems Requirements, McDonnell Douglas Astronautics Company

I. J. Webster Mission Integration/Launch, PAM Programs, McDonnell Douglas Astronautics Company

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Scholarly Commons Citation Bryan, C. E. and Webster, I. J., "Processing and Deploying the McDonnell Douglas Payload Assist Module (PAM)" (1983). The Space Congress® Proceedings. 7. https://commons.erau.edu/space-congress-proceedings/proceedings-1983-20th/session-iib/7

This Event is brought to you for free and open access by the Conferences at Scholarly Commons. It has been accepted for inclusion in The Space Congress® Proceedings by an authorized administrator of Scholarly Commons. For more information, please contact [email protected]. PROCESSING AND DEPLOYING THE McDONNELL DOUGLAS PAYLOAD ASSIST MODULE (PAH) C. E. Bryan, Unit Chief Electronic Systems Requirements McDonnell Douglas Astronautics Company I. J. Webster, Manager Mission Integration/Launch PAM Programs McDonnell Douglas Astronautics Company

ABSTRACT able vehicle/spacecraft, and the thermal protection for the entire cargo element. The This pap*r presents the flow of the opera­ PAM ASE also provides the ability, should it tional PAM system from the time processing is be necessary, to safely abort a deployment started at the launch site through deployment attempt and return the cargo element to the from the Orbiters. It addresses the ground landing site. The need to provide ASE opera­ checkout activities, in-orbit operations tional failure tolerance and to meet the including crew and ground personnel system Orbiter payload safety requirements resulted evaluation and command activities, and PAM in significant mechanical, ordnance, and deployment from the Orbiter. Additionally, avionics redundancy, all of which bear transfer orbit errors for two PAMs used on directly on checkout of the ASE. STS-5 are presented. The avionics, (Figure 2) performs the The PAM ground processing approach affords following functions: maximum assurance of a flight-ready PAM prior 1. Receives and implements commands from to mating the spacecraft and provides a cargo the Orbiter general purpose computer (GPC), element that is fully verified as flight- including keyboard entries, and from the ready before integration with the cargo Orbiter standard switch panel. integration test equipment (CITE) and the 2. Conditions Orbiter power and distri­ Orbiter. The PAM system design and on-board butes it to the PAM and spacecraft. data displays give the astronauts the capa­ 3. Provides closed-loop monitoring and to the bility to evaluate the status of the PAM's sequencing of PAM systems subservient health and deploy the PAM/spacecraft without crew commands. air-to-ground data or communications. 4. Generates system status information for display to the crew and down!isting. SYSTEM DESCRIPTION 5. Processes and provides data for recording on the Orbiter payload data The PAM-D is a system to deliver spacecraft recorder. from the STS Orbiter low earth orbit to usable higher energy orbits using a spin The sequence control assembly (SCA) is the stab!ized solid propellant stage or expend­ microprocessor-based master controller and able vehicle (EV). The EV and spacecraft are data acquisition system for the PAM. Moni­ supported vertically in the Orbiter bay by toring and closed-loop sequencing of the the PAM airborne support equipment (ASE). interfacing PAM ASE and vehicle systems are All avionics and electro/mechanical hardware provided. SCA interfaces with the systems , needed for on-orbit system evaluation, EV and include discrete outputs, discrete and analog spacecraft spin-up and deployment and PAM EV inputs, programmable pulsed outputs, and flight are provided. The PAM system elements programmable input event counting or time are shown in Figure 1. interval measurements. The SCA also inter­ faces with the Orbiter data bus for command and data communications with the Orbiter GPC Airborne Support Equipment Configuration and the crew keyboard/CRT. Two SCAs are The ASE consists of the cradle assembly and used: one as the primary SCA and the other as the spin, separation, thermal protection, and the backup. avionics systems. The ASE provides the structural and electrical interfaces with The spin system distribution box (SSDB) both the Orbiter and the expendable vehicle, provides for control, fusing, and distribu­ the spin and deploy functions for the expend­ tion of the Orbiter payload bus 28-VDC power

I IB-8 PAF With Equipment Installed

SRM

Spin Table and Separation Spacecraft System Sun Shield ASE Avionics

Orbiter Longeron Fittings (4)

ASE Thermal System

Cradle Assy Orbiter Keel Fitting

Figure 1. STS PAM-D Configuration

Vehicle J Slip Rings

Commands ASE System PAM-D ASE Avionics • Spin System Monitors 2 — Sequence Control • Restraints Assemblies (SCA) • Heaters Power 1 — Signal Conditioning • Sunshield Unit (SCU) 1 — Spin System Distribution Box (SSDB) 1 - PCM Encoder

ASE

Figure 2. PAM-D Avionics

IIB-9 to the RAM and spacecraft systems. In addi­ The PAM-D hardware flow from factory through tion, it provides relays and control circuits launch and return is shown in Figure 5. PAM for the spin motor drive, sunshield, system preparation starts in the Explosive restraints, heaters, index solenoid, and spin Safe Area 60 (ESA-60), where the ASE is brake systems. refurbished and retested, and the expendable vehicle hardware is assembled, tested, and The signal conditioning unit (SCU) includes mated with the ASE. Expendable vehicle spin command and monitor circuitry for the S&A balancing is done in the Delta Spin Test devices, deployment bolt cutters, SCA enable Facility (DSTF). The spacecraft is mated to function, abort function, and the predeploy- the PAM in the north bay of ESA-60 and the ment functions on the expendable vehicle. interfaces with the PAM are verified. The Signal conditioning is provided for compati­ complete payload element is transported to bility between source signals and receiving the Vertical Processing Facility (VPF) using units. The SCU also contains isolated power the Pam transporter. The PAM and spacecraft supplies, which provide the required regu­ interfaces with the Orbiter are verified by lated power for the PAN systems. NASA in the VPF with the cargo integration test equipment (CITE) before the PAM-D/ Expendable Vehicle Configuration spacecraft are moved to the Rotating Service The expendable vehicle hardware (Figure 3) Structure (RSS) and installed in the consists of the Thiokol Star 48 motor, the Orbiter. Once in the Orbiter, the actual PAM PAF structure, the spacecraft separation and spacecraft to Orbiter interfaces are hardware, the motor safe and ami (S&A) verified. The PAM-D airborne support device, and the electronic sequencing sys­ equipment is returned in the Orbiter to KSC tem. The fully redundant electronic and then returned to MDAC at the Orbiter sequencing system (Figure 4) consists of a Processing Facility (OPF). timer assembly, an electronic control assembly (ECA), and batteries. Expendable Vehicle Processing Final assembly of the PAM expendable vehicle PAM-D LAUNCH SITE PROCESSING hardware is done at the launch site. Compo­ nents delivered to the launch site are: The PAN hardware and GSE design provide • Payload attach fitting complete with launch site checkout capability of the PAN all electrical hardware and functionally systems, including redundant subsystems and tested with the spacecraft. components, with a minimum number of electri­ • Solid rocket motor witti flame sticks cal disconnects. This is accomplished by the and TBI's installed. use of GSE and test connectors on the PAM • Loose items, including safe and arm hardware that were designed specifically for (S&A) assembly, confined detonating fuse that purpose. The PAM test philosophy is to (CDF), tumble system, brackets, etc. functionally verify the entire PAM system at the launch site, with minimum disruption of The expendable vehicle buildup and checkout assembled hardware. is designed and sequenced to ensure that the

Cradle Reaction Fitting Cradle Reaction Fitting Separation Spring (4) Cradle Support Base Brake Separation Band Electrical Spacecraft Slip-Ring Electrical Interface Spacecraft Separation Band Clamp

Spacecraft Electrical Interface s Index Solenoid (2)

Spacecraft Separation Spring (4) Payload Attach SRM Motor Fitting Spin Bearing Cradle Reaction Fitting Cradle Reaction Fitting Telemetry Antenna Tumble Installation Spin Table Drive Motor (2) Figure 3, STS PAM-D inboard Profile

IIB-10 PAM-D PAM-D Vehicle Spacecraft ASE

Redundant Redundant Batteries Timers

Redundant Safe and SRM Separation Arm Device Igniter Switches Electrical Control Assembly Spacecraft External (Voting Separation Power/Commands Logic) Ordnance

Function Monitors Tumble System Ordnance Safety Critical Commands

Safety Critical Function Monitors

Redundant Safe Spacecraft and Arm Position Commands Inputs

Safe and Arm Position Monitors Spacecraft Monitors

Figure 4. STS PAM-D Sequencing System Functional Diagram expendable vehicle is flight ready before it the expendable vehicle system test. This is mated to the ASE. test verifies that the interfaces to the spacecraft and to the ASE are correct, that Expendable Vehicle Component. The solid the flight sequence functions are correct, rocket motor (SRM) is inspected, cold soaked, that the necessary redundancies are In and X-rayed before it is moved to ESA-60 that all ordnance firing circuits are of where it is leak checked and assembled into stray voltage, and that they function satis­ the expendable vehicle. The SRM S&A and RAM factorily under simulated ordnance loads* separation ordnance devices are subjected to This test is depicted in Figure 6. the following testing prior to installation: • S&A electro/mechanical checkout ASE Processing • Spacecraft and motor deploy cutter ASE processing starts with the and electrical checkout refurbishment of the ASE from a • Yo-weight deploy cutter electrical mission. Following refurblshMent* checkout conducted to confirm that the ASE Is tioning properly. The PAF is received at ESA-60 and the S&A, test batteries and CDF are installed. Func­ ASE Post Fl1ght Refurbl tional testing of the PAF is accomplished of the ASE In ESA-60, It is the after assembly of the expendable vehicle is sunshield and thermal arc complete. and the cradle Is for the sion* The Expendable Vehicle Buildup and System Test. bolt cutters, 'and After the expendable vehicle components are Refurbishment of ASE hts » processed, they are assembled to provide the open anomalies its PAN expendable vehicle. The expendable does not vehicle is weighed and then transferred to operational -Iffe Is ta tit the DSTF for spin balancing, after which it correction of as t is retruned to ESA-60 and electrically con­ of a of tlit nected to the PAM system test equipment for is, tii tit?

111-11 MDAC, Huntington Beach Thiokol, Elkton Spacecraft Contractor • Stage • ASE • SRM • Spacecraft Hardware • GSE

ESA-60 DSTF • PAM-D Preparation • Expendable • Spacecraft/PAM-D Vehicle Spin (Payload) Mate Balance • Interface Verification

VPF • Spacecraft/PAM-D CITE Verification

I

' STS PAM-D RSS/Pad Mission ———> • Orbiter/Payload Mate • Integration/Verification Orbital • Closeout Store V______J Operations

OPF • Remove STS STS PAM-D ASE PAM-D ASE

Figure 5. STS PAM-D Program Flow thermal blankets, ordnance-limiting resis­ tors, and adjustment of the limit switches RAM System Checkout during preparations for normal checkout. Set Control Console Resolution of anomalies is by standard material review action. Any components approaching their expected life limit are removed and replaced. The removed units are returned to MDAC HB or to the supplier for PAM System Checkout refurbishment. Set Console

Cradle Systems Testing. The PAM-D ASE, including redundant systems, is functionally verified in preparation for each mission using the PAM system test equipment. This Connect Test Cables Battery and Ordnance test equipment also simulates all interfaces Simulator, Control Monitor with the Orbiter and verifies these inter­ Install Test Batteries faces to be correct and functionally satis­ Conduct Expendable Vehicle Systems Test factory as part of the testing. The testing — Separation Switches is accomplished in the following major steps, — Redundant Circuits 7. — Timer Systems which are illustrated in Figure Verify Ordnance Circuit Loads * Pre-test Preparations — A PAF simula­ Perform Flight Sequence tor is installed to functionally simulate PAF Verify Spacecraft Interface Circuits interfaces for test. Addi­ Disconnect Test Battery Connectors and spacecraft tionally, a moment of inertia simulator is attached to the spin table to provide moment Figure 6. Expendable Vehicle Systems Test — ESA60 of inertia for spin operations.

IIB-12 • Fault Simulation Test - The PAF inter­ Pay load Attach RAM System Checkout Fitting Simulator Set Control Console faces are faulted by PAM 500 software com­ Mass Simulator mands to verify that the SCA will recognize, correctly respond to, and report conditions such as low PAF battery voltages, incorrect RAM System Check­ out Set Console S&A talkbacks, malfunctioned separation switches, and other safety monitor conditions. • SCA Control Transfer Test — This test checks the capability to switch from the pri­ mary SCA to the backup SCA and to continue a *** Strain Indicator mechanical sequence or stow operation. System Install Mass Simulator Install and Connect RAF Simulator Expendable Vehicle to ASE Mate Connect Test Cables After the expendable vehicle and ASE testing Pre-Power Test - Resistance and Isolation are completed, the expendable vehicle is Measurements positioned on the spin table and mated to the - Spacecraft Continuity Test ASE structurally and electrically. Measure­ - VCO Calibration ments are made to verify proper alignment of Perform Cradle Systems Testing the expendable vehicle and the cradle - Verify PCM Interfaces trunnions. - Perform SCA Functional Test - Perform Sunshield and Restraints Operation Test PAM System Test - Verify Spin Control System The PAM system test functionally tests the - Verify Deployment Sequence assembled expendable vehicle and ASE, again - Verify Stow Sequence using the PAM system test equipment to simu­ - Verify Redundant Logic - Fault Simulations I late the Orbiter interfaces (Figure 8). The mm mmmmmmmmmMmmammmmmmmmmmmmmwmmmmmmmmmwd weight of the expendable vehicle is lifted to Figure 7. Cradle Systems Testing — ESA60 simulate the in-orbit zero-g environment, and the PAM electrical configuration is as near The Isolation and continuity of power the in-orbit configuration as possible. system is checked, The Orbiter, SSP and T-0 power PAM system test verifies that the deployment supply set-ups are accomplished and verified and spacecraft to umbilical and SSP interface wiring is verified. • Cradle Subsystem Verification -- The Spacecraft PAM System Checkout following subsystems are verified as Set Control Console indicated: Simulator -- Power supply systems in the SSDB SCU and SCA's are functionally tested PAM System -- PCM system is activated and the FM Checkout Set YCO's are adjusted Console - SCA functional tests verify data bus interfaces, command and monitor modes, SCA self test, and invalid command response - Spacecraft interface circuitry is tested to verify SCA to spacecraft discrete commands, discrete monitors, and analog input circuits. Strain Indicator System • Cradle Software Verification — This system test accomplishes an SCA RAM and ROM Verify Cradle/PAF Interface Circuits — Connect Test Cables memory dump and automatically compares that — Spacecraft Interface Circuit data with the defined SCA software Continuity and Isolation Test configuration. — Restraint Alignment • Cradle Flight Sequence Test ~ Func­ — SSP Control Test — Rotate S&A tional verification of the mechanical Verify Spacecraft Interface sequences at mission level rpm, verification — Connect Spacecraft Mission Peculiar of redundant power system operation, and Harnesses ability of system to operate with a single — Connect Spacecraft ASE power bus during sunshield, restraints and — Install and Connect Spacecraft Simulator — Perform Spacecraft Interface Test spin operations. The following operational Conduct Integrated Systems Test systems are verified via SCA software. — Verify Deployment Sequence - Sunshield motors and limit switches — Verify Flight Sequence Timing - Restraint motors and limit switches — Verify Stow Sequence System -- Spin system control — Remove Hoist Plate and GSE Clamp Band -- Terminal deploy sequence -- Stow sequence Figure 8. PAM Systems Test - ESA60

IIB-13 and stow sequences operate properly in their vertical payload handling device (VPHD) with operational configuration and that any space­ other cargo elements for the mission. PAN craft ASE needed to interface the spacecraft testing in the Cargo Integration Test Equip­ to the RAM functions satisfactorily when ment (CITE) demonstrates compatibility integrated into the system. The RAM system between the RAM cargo element and the Orbiter test is the final test to verify that the RAM and between the RAM cargo element and other will successfully accomplish the in-orbit cargo elements. deployment sequence and that the system is ready to receive the spacecraft. VPF and RSS Handling Operations The RAM cargo element is delivered to the YPF Spacecraft Hate and Checkout in the MDAC handling frame, which also serves Spacecraft mating and checkout are shown in as the transporter. After the necessary Figure 9. The spacecraft is positioned on cleaning of the transporter's external sur­ the payload attach fitting and the spacecraft faces at the YPF airlock, the transporter, separation clamp assembly is installed at the with the cargo element, is moved into the RAF separation plane. An interface verifica­ high bay and the transporter's environmntal tion test is then performed to verify that cover is removed. The cargo element is the circuits between the spacecraft and the rotated from the vertical to the horizontal RAM are functional. Power-on and power-off (liftoff) orientation to allow hoisting and stray voltage testing is accomplished, positioning into the VPHD. The hoist beams yo-weight explosive cutters are connected, are attached to the cradle and the cargo and the thermal control system is installed. element is hoisted into its cargo location. Sunshield and heater circuit verification Access platforms are positioned and the tests are the final RAM tests to be conducted interfacing cables are connected to the RAM before the cargo element is placed in the for interface testing. transporter and delivered to the VPF. When the testing is complete, the entire VPF AND IN-ORBITER OPERATIONS cargo is prepared for transfer to the payload cannister and transportation to the Rotating The RAM/spacecraft assembly (RAM cargo ele­ Service Structure (RSS). The payload can­ ment) is delivered to the Vertical Processing nister and cargo are transported to the RSS Facility (VPF) where it is placed in the by NASA, where the cargo is placed in the payload ground handling mechanism (PGHM) inside the RSS. When the cargo is ready to be installed in the Orbiter, the PGHM is moved to the orbiter and the cargo is trans­ RAM System Checkout Set Control Console ferred to the Oribter. CITE Testing The RAM ASE provides the interface between the RAM cargo element and the orbiter. Since this ASE has just been returned from a pre­ vious mission, these interfaces are already verified to be correct unless there have been changes to either the RAM or Orbiter. The functional tests presently being accomp­ lished in the CITE verify the functional compatibility of the RAM hardware interfaces RAM System Checkout with the CITE's simulation of the Orbiter. Set Console Also, cargo-element-to-cargo-element compati­ Strain Indicator bility is verified. System f Prepare for Spacecraft Mate The test sequence consists of powering up the « Separation Springs RAM and interfacing simulated Orbiter equip­ t Weights i* ment to: « Spacecraft Clamp • Verify the four on-board displays and a * f>AM-D/$pacecraft Interface Verification sample of the ground displays with the SCA in a Connection an idle state. — On/Off Stray Voltage • Verify that the GPC can send commands — Cutter Connection — 'Vp Connection to the SCA and that the correct response is * returned to the GPC. * Tifee • Verify the correct operation of all '» switches on the SSP. i* Circuit Verification Test • Conduct an abbreviated mission sequence to verify correct commands and responses per §. — the deployment test scenario.

1IB-14 • Verify the RAM PCM data by recording the astronauts via the keyboard of the multi­ the data during the above tests. function CRT display system (NCOS). For a standard PAM mission, the following data bus After completing the functional test, the commands are required to operate the PAM. ability to charge the RAM batteries through the T-0 umbilical is verified and stray- Crew keyboard initiated voltage measurements on the RAM ordnance • Close sunshield lines are made. • Open sunshield • Start mechanical sequence Orbiter To RAM Testing After the RAM is installed in the Orbiter, GPC program initiated the hardware interface test is repeated. The • Start terminal deployment sequence test is the same as conducted in the VPF. • Deployment fire Power-on and power-off checks for stray voltage are conducted, and final ordnance Additional mission-unique keyboard commands connections are made. The cargo element is are used to control the spacecraft equipment then closed out, the S&A mechanical safing via the data bus and SCA. Backup commands pin removed, and the sunshield secured. The for some PAM systems are also provided. PGHM is then retracted into the RSS. The Orbiter bay doors are closed. The only PAM and spacecraft data are displayed to the functional operation prior to launch is crew on the MCDS CRT. For each PAM/ trickle charging of the PAF batteries up to spacecraft there are four CRT display pages. launch minus 1.5 hours. The RAM-related fields of the displays are standardized, whereas the spacecraft fields PAM-D IN-ORBIT OPERATIONS vary from mission to mission. Figures 11 and 12 show two of the displays used for the The RAM system and its interfaces with the ANIK-C mission. The deployment display is Orbiter are designed to minimize the crew designed to facilitate the monitoring and activity required for deployment while pro­ commanding of the time-critical deployment viding on-board visibility of RAM health and operations phase. The control display pro­ safety status. The on-board command and data vides monitor and command capability used display capability allows the astronauts to primarily during the other PAM activity deploy the RAM/spacecraft in the absence of phases. air-to-ground data or communications. All PAM data and most spacecraft data dis­ For Orbiter launch and ascent, the PAN system played on the CRT are transmitted to the GPC is configured with the sunshield open, pay- from the SCA via the payload data bus. On load restraints inserted, and avionic systems some missions, spacecraft data that Is Input powered down. The on-board RAM activity is to the payload data Inter! eaver (PDI) is also conducted in four major phases: sunshield used to drive the CRT displays. closing (post-insertion); pre-deployment checkout; deployment operations; and sun- The data bus data and PDI data are downlinked shield opening (pre-deorbit). for ground display and e'valluation at the JSC Mission Control Center and! the spacecraft- Command and Monitor Interfaces provided Payload Operations Control Center. me Orbiter crew commands and monitors the PAM and spacecraft special Ists, representing RAM and spacecraft systems using standard the- customer and his contractors, evaluate Orbiter equipment. Commands from the Orbiter the data with the assltance of HftSA to RAM are sent from the standard switch operations personnel • panel (SSP) or the general purpose computer (GPC) via the payload data buses. Su nshl ell d Cl psl ng { Pos 'After the Orbiter has aclfcwel The SSP, installed in the aft flight deck, is tlon and the payload ban" doors ham used for control of power on/off, required opened, the PAM sunshield oust be closed aiiii safety functions, and selected backup com­ the heaters must be enabled in trier to mands. The normal layout of the SSP is shown thermal If protect the Pill and spacecraft* in Figure 10. Each of the twelve switches The requl red sequence 1 icflll udes: has an associated talkback indicator. Power Power up MM control functions include SCA power, 1 sol /PCM PAN power, SCA enable, and ASE/PAM heaters. Goraand ""ell ose sunsMel dP w it keyboard Safety related functions include RAM S&A arm, SCA, automatical If ell ose sunsMel d spacecraft S&A arm, and deployment pre-arm. Power tip MM heaters Backup commands include vehicle ordnance pre­ Power down PAN SCfts arm, deployment arm, deployment fire, and stow sequence initiate. Commands to the PAM The thenaostatlcal If control 1 od to^ 1 n from the GPC via the payload data buses may conjunction with the sunsMol d, control the be initiated by pre-stored GPC programs or by PAM and spacecraft temperatures wt thin

IIB-15 Dp-Arm

DN-Safe

Up Up

(DS1) (DS2) (DS3) (DS4)

S/C Deploy Stow S&A Pre Arm Seq Arm Pre Arm Initiate

1 (S5) — •(S1) ^- (S2) (S3) Safe A Safe A Off Safe A

SWPwr Not Used

MCR2) (DS8) (DS9) (DS10) (DS11) (DS12)

ISOL/PCM ASE/PAM SCA Pwr Heaters Enable On Auto SCA-1

(S9) - (S8) Off Off A

Figure 10. Standard Switch Panel acceptable limits during the on-orbit opera­ PAM and spacecraft data is performed by the tions preceding deployment. crew and, when air-to-ground communications permit, by ground personnel. The required Predeploy Checkout Phase sequence includes: Prior to entering the predeploy PAM checkout Power up PAM SCA's procedure, the updated deployment time and Verify GPC to SCA 2 communications attitude are uplinked to the Orbiter. The Verify safety status crew enters this updated deployment time via Verify GPC to SCA 1 communication the table maintenance display. The PAH Verify safety and health status from sequence control process, which resides in all display pages. the GPC software, is used to communicate with the PAM SCA. The crew starts this process and initiates the deploy countdown by a key­ Deployment Operations Phase board entry via the deployment display. In The Deployment Operations phase includes the addition, the crew enters the updated deploy­ final sequence of activities necessary to ment attitude into the Orbiter GNC system. deploy the PAM/Spacecraft from the Orbiter. A summary of these activities is shown in The predeployment checkout procedure is Table 1. entered approximately one hour before the • Orbiter Maneuver -- The crew commands planned deployment time. The objective is to the Orbiter to maneuver to the deployment verify that the PAM and spacecraft systems attitude. This pre-stored sequence is are in the safe and flightworthy state initiated at deployment minus 40 minutes by required for initiating the deployment opera­ keyboard entry via the universal pointing tions phase. A comprehensive review of the display.

IIB-16 C XXXX/201 ANIK-C DEPLOY XX X DDD/HH : MM « SS^ SCA CLK XX XXXX DDD/HH:MM:SS DEPLOY X XXX/XX:XX:XX ITM EVENT SCHD CPLT 1 COUNTDOWN XX i XX I XX XXXX 2 MECH SEQ START XX Ml CUR M2 CUR STBD RSTRNT OUT XS X.XX X.XX PORT RSTRNT OUT XS X.XX X.XX SUNSHIELD OPEN XS x.xx x.xx SPIN XXX. XS XXX. X X XX. X XX. X 5/6 S/C ASE X/ENCDR ON XSXS 7 CONFIG XS 8 I NT PWR XS XXX XXX XXXX TERM SEQ START 3100 XS VEH ORD PREARM 1S30 X ABORT SEQ SSP PAM 58 A XS SSP VEH ORD SF X SSP S/C 58 A xss SSP DPL PA SF X SSP DEPLOY PREARM >:30 X SSP PAM SI A SF X 9 S/C SIU DIS :25 XS 10 S/C SIU ENA X DEPLOY ARM 505 X SSP S8A SF X DEPLOY FIRE «00 X 11 DECONF X 3 DEPLOY INHIBIT XXX SSP STOW SEQ X 4 RESET (XX) k j

Figure 11. ANIK-C Deployment Display /^XXXX/211 ANIK-C CONTROL XX >( DDD/HH:MM«SS ^ COUNTDOWN XX J XX J XX SCA CLK XX xxx>( DDD/HH:MM:SS PR I SCA X XXXX DEF>LOY > C XXX/XX:XX«XX I/O INIT 1 DEPLOY ORD ARM SUNSHIELD SLF TEST XXXX PREARM XSXS CLS SSHLD 5X5 OUTPUTS XXX ARM CMD XSXS OPN SSHLD 6X5 STATUS XXXX ARM BUS XSXS MTR 1 2 SPIN RATE FIRE CMD XSXS ENA 7X 8X LIM ORIDE 2X FIRE RLY xsxs DIS 9X 10X SLOW 3X VEH ORD OPN 11X 12X NORMAL 4X PREARM BUS xsxs CLS 13X 14X S/C SYSTEMS 1 2 ARM BUS xsxs ON 15X 16X 2-1/22 ASE ON XSXS PAM StA XXX OFF 17X 18X 23 OFF SAFE XXXS CUR X.XX X.XX 24/25 BUS OFF XSXS S/C ORD 1 2 POS IXXXX XXXX 26/27 TCHG ON XSXS S/C StA xsxs 2XXXX XXXX 28/29 OFF XSXS 34 SAFEXSXS 3XXXX XXXX SIU ON XSXS SEP SW CL xsxs 30/31 ENA XSXS AKM ENA xsxs 32/33 DIS XSXS INH SW xsxs RLY BUS ON PCM CLK ON XSXS LVLV OPN 1 xsxs IS 25 35 45 ODD STAT GO XSXS 2 xsxs PAM SEP SW CL EVEN STAT GO XSXS OMNI ENA 1 xsxs BV AS B5 BATT T GO XSXS 2 xsxs 16V AS BS (XX) 1

Figure 12. ANIK-C Control Display

• Mechanical Sequence — This sequence The mechanical sequence is initiated by a provides for the accomplishment of three crew keyboard command via the deployment major operations required to configure the display. Upon receipt of the command, the PAM system for deployment: sunshield PAM SCA automatially controls the sunshield, opening, payload restraints withdrawal, and restraints and spin-up operations. The SCA spin-up. In addition, the seqeunce includes initiates withdrawal of restraints only after power-up of the spacecraft for deployment. verifying that the sunshield is open and

IIB-17 Table 1

RAM Deployment Operations Phase

Tine Start Norn Duration Event (min:sec) (min:sec) Commanded By

Maneuver to deploy attitude -40:00 12:00 Crew (KYBD)

Start mech sequence -15:00 Crew (KYBD) Open sunshield -14:50 0.23 SCA

Right restr withdrawal Sunshield open 0:50 SCA

Left restr withdrawal Right restr withdrawal 0:50 SCA

Spin up to deploy RPM Left restr withdrawal 0:30 SCA

Spacecraft ASE pwr turn-on -12:00 Crew (KYBD)

Spacecraft configuration Crew (KYBD) Start terminal deploy -3:00 GPC sequence

Yeh ord prearm -1:30 SCA

RAM S&A arm Crew (SSP)

Spacecraft S&A arm Crew (SSP)

Deployment prearm Crew (SSP)

Deployment arm -0:05 SCA

Deployment fire 0:00 GPC

Close sunshield +0:30 0:23 SCA

Enable post-deploy heaters Crew (SSP)

SCA power off Crew (SSP)

Maneuver to seperation +3:00 Crew (KYBD) burn attitude

OMS seperation burn +15:00 Crew (KYBD)

Maneuver to window +29:00 Crew (KYBD) protection attitude

initiates spin-up only after verifying that The terminal deployment sequence, exercised the restraints are withdrawn. After the com­ by the RAM SCA, is automatically initiated by pletion of spin-up, the system is maintained a GPC to SCA data bus command at 3 minutes at the deployment spin rate (typically 50 before deployment. At deployment minus 90 rpm) until deployment. seconds, the SCA issues the vehicle ordnance • Terminal Deployment and Sequence — pre-arm command to the RAM vehicle. The crew This sequence includes the RAM and spacecraft then activates and verifies three ordnance ordnance system arming and the actual functions from the standard switch panel: deployment of the RAM and spacecraft from the RAM S&A arm, spacecraft S&A arm, and deploy­ Orbiter. ment pre-arm.

IIB-18 Any required final spacecraft configuration springs, achieving a velocity of approxi­ commands are issued by the crew from the mately 2.5 feet per second. keyboard. The pilot makes final verification of Orbiter attitude and rates. Figure 13 shows a photograph of the RAM and SBS-C spacecraft being deployed from the Orbiter on the STS-5 mission At deployment minus 5 seconds, the SCA makes a final verification that the PAN vehicle • Postdeployment Securing -- After the systems and the spin rate are in a go RAM has cleared the Orbiter, the SCA auto­ condition. If the state is no-go, the SCA matically closes the sunshield to protect the automatically halts the deployment sequence. ASE for the remainder of the Orbiter mis­ If the state is go, the SCA commands sion. The crew turns on the RAM cradle deployment am. At the stored deployment heaters and then powers down the spacecraft time, the GPC then automatically issues the and RAM ASE. deployment fire command to the SCA, which activates the fire relays and the deployment In order to minimize Orbiter contamination bolt cutters are fired. from the plume, the crew commands the Orbiter to perform a series of maneuvers to place the The spinning RAM/spacecraft is propelled from Orbiter in a protected attitude and at a safe the spin table by the four separation distance from the RAM at motor ignition.

Figure 13. PAM/SBS-C Deployment from Columbia on 11/12112

IIB-19 Sunshield Opening (Predeorbit) errors it is clear that Orbiter and PAM Prior to closlg the Orbiter payload bay doors pointing error were each very small and that in preparation for deorbit, the RAM sunshield the PAM propulsive performance was very close is opened to ensure proper dynamic clearance to predicted. with the doors. The required sequence includes: Table 2 • Power up PAM SCA's Performance Error • Command "open sunshield 11 via keyboard (STS-5 Mission Performance Reconstruction) • SCA automatically open sunshield • Power down PAM heaters Error = Actual - Predicted • Power down PAM SCA's Parameter BS-C ANIK-C 3

STS-5 PAM MISSION PERFORMANCE • Velocity (fps) 16.7 14.0 ~ *sp (sec) 0.56 0.48 The spacecraft flown on STS-5 (SBS-C and Apogee Altitude* (nmi) 148.1 232.3 ANIK-C3) are both in stationary orbit and are Perigee Altitude (nmi) +0.1 -0.1 performing very well. The entire STS-5 mis­ Inclination (deg) -0.02 0.04 sion was excellent, with the crew deploying Augment of perigee -0.17 0.11 both PAM's flawlessly. The PAM ASE and Spin rate 0.4 0.6 expendable vehicle performance were anomaly free. The transfer orbits provided were * Based on integrated value at first apogee accurate, as indicated by the performance error summary in Table 2. From these total

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