[american institute of aeronautics and astronautics 31st joint propulsion conference and exhibit -...

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AlAA 95-2409 Peacekeeper Flight Termination Ordnance System James A. Ramsey, Major USAF, Space and Missile Systems Center San Bernardino, CA 92408-1621

Donald R. Sessler and Lien C. Yang TRW Strategic Systems Division San Bernardino, CA 92402-1310

31 st AIANASMEISAUASEE Joint Propulsion Conference and Exhibit

July 10-12,1995/San Diego, CA For permission to copy or republish, contact the American inatkute of Aeronautic8 and Astro~utlcs 370 L'Enfant Promenade, S.W., Washington, D.C. 20024

PEACEKEEPER FLIGHT TERMINATION ORDNANCE SYSTEM

James A. Ramsey, Major WAF, Space and Missile Systems Center

San Bernardino, CA 92408-1621

Donald R. Sessler and Lien C. Yang TRW Strategic Systems Division

San Bernardino, CA 92402

Abstract tion. For the Post Boost Vehicle (PBV), redundant Through-Bulkhead Initiators (TBIs) actuate a Fuel Line Severance Assembly (FXSA) to shut off the liquid propellant supply to the enginehozzle assembly and vent the fuel from the storage tank. This assembly can also be actuated to dispose of excess fuel after normal PBV operation. These explosive components are initiated by Exploding Bridgewire Initiators (EBWIs) which are integrated in a high voltage capacitive discharge type Firing Unit 0. Eight FU/EBWI assemblies (two per stage) are used in each Peacekeeper flight test. The detonation outputs of the EBWIs are coupled to the LSCAs and TBIs by flexible Ordnance Transmission Assemblies (WAS), which are made of Confined Detonating Cord (CDC) and booster charge end tips. The configurations are illustrated in Figure 1.

The high voltage capacitor in each FU is charged to and maintained at 2300-3200 V before the flight by the Fu's DC-DC converter. The discharge into the EBWI is controlled by a spark gap vacuum trigger switch assembly. The triggering is initiated by a high voltage pulse generated by discharge of a low voltage capacitor into a step- up transformer.

The FU also contains a logic circuit to execute Command Destruct (CD), destruct due to Premature Stage Separa- tion (PSS), or an inhibit function to allow the override of the FU f ~ n g in a planned staging event. The high voltage capacitor voltage, mgger capacitor voltage, and the inhibit signal are continuously monitored before and during the flight via telemetry.

A computer based FU Test Set (nrrS) is used to perform detailed FU preflight checkout at Vandenberg AFB. This self-calibrated equipment measures FU inpufloutput resistances, input signal currents, reverse cumnts, monitor voltages, and destruct output current into an EBWI simulator, at input voltages of 24 and 32 Vdc. It provides verification of inhibit, CD, and PSS functions and indicates that all parameters are in agreement with the limits used in the acceptance test at the factory.

The Peacekeeper Intercontinental Ballistic Missile flight tests are equipped with a highly reliable Flight Termina- tion Ordnance System which is capable of terminating the flight from a ground command or in the event of premature missile breakup. The system is based on an exploding bridgewire initiator and a high voltage fuing unit. It has more inherent safety features as compared to con- ventional hot-bridgewire-detonator and safe-and-arm device based destruct systems. The detonation output of the exploding bridgewire initiator is transmitted by Ord- nance Transmission Assemblies to Linear Shaped Charge Assemblies and Through-Bulkhead Initiators for motor case and fuel line severances. This paper repom the high- lights of the design, performance, and testing of this system and its components.

4 I. Introduction

In the flight tests of the Peacekeeper Intercontinental Bal- listic Missile System, both in earlier Development Test and Evaluation (DT&E) and later Operational Test and Evaluation (OT&E), the ability to reliably thrust terminate an errant missile is required. This termination may be initiated by a ground command from range safety to abort the flight in progress or in an automated manner if a missile breakup occurs. This capability is provided by a Flight Termination Ordnance System (FTOS).

€TOS is a subsystem of the Instrumentation and Flight Safety System (IFSS). IFSS provides telemetry for instrumentation data, antennae and receivers for respond- ing to ground commands, and electrical power and control to the FTOS.

The fundamental requirement of FTOS is to render all missile propulsion components non-propulsive. For the three solid rocket booster stages, redundant Linear Shaped Charge Assemblies (LSCAs) are designed to sever the Kevlar composite based rocket motor cases upon activa-

d This paper is a work of fhe U.S. Government and is not subject to copyright protection in the United Stares

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Figure 1. Peacekeeper lTOS System Layout

To date the FTOS has been successfully deployed in 36 Peacekeeper flight tests. An aging and surveillance program on FTOS components is under way to gather data for trend analysis life estimates. It is envisioned that the FTOS can be easily adapted for other aerospace applica- tions such as launch vehicles.

This paper will highlight the design features and performance of the FTOS and its components. In addition, the design philosophy, development history, qualification and acceptance testing, which are the key elements of the system's success, will be described.

II. Svstem Desien

The design is fairly simple, as illustrated in Figures 1 and 2.

Redundancv

The FTOS system reliability requirement 0.99999 can only be achieved by using a redundant design. The de struct components including the FSS batteries, FUs, Com- mand Receiver Decoders (CRDs), accelerometers for inhibiting FUs, LSCAs, Pressure Switches (PS), and OTAs for F'LSAs are implemented in separate hardware chan- nelsA and B. The FLSA is a single device with identical dual design in explosive train and mechanical parts. The Arm/Disatm ( A D ) switch is a single package with isolated switch wafers.

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Figure 2. Peacekeeper FTOS System Schematic

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Localized Staee Destruct Components

The destruct components for each stage, including the FUs, are located on the same stage. This implementation provides a level of autonomy of FUs; i.e., FUs are capable of executing the destruct after the electrical power has been cut off from FUs during a premature stage separation event. It also offers good system weighuspace sav- ings.

Premature Staee Separation Signal

The IFSS provides a signal ground from the PBV on two wires for each channel in the main IFSS cable in the raceway passing through the mechanical staging connectors. Both wires are connected through each Fu. Loss of this connection (>25 kohms impedance to the ground) continuously for 2.5 ms is the PSS signal. For any FU to execute the destruct, loss of both PSS signals to that FU is required.

AcceleratiodPressure Switch Inhibit (AIPSI) Sienal

The PSS signal is lost during normal staging events. The inhibiting of the destruct under this condition is mecha- nized through inhibit accelerometers in IFSS and redundant pressure switches on each stage. The accelerometer senses the decrease of the missile acceleration during the tailoff of each of the three solid stages. At a preset low acceleration level, IFSS reverses polarity on a low voltage (1.6 to 5.0 Vdc) signal downstage to the departing stage FUs via a PS. The PSs are normally open and actuated by the explosion pressure of the staging ordnance (separation rings) on which they are plumbed. The inhibit signal is crossover linked between the A and B chm- nelson each stage, Le., both FUs receives the inhibit signal from bath PSs. Only one inhibit signal is required to inhibit the FU. Alps1 is not implemented for PBV or shroud separation because any PSS signal separations would be abnormal.

Command Destruct Sienal

All FUs per channel are connected in parallel by a pair of wires to the FSS CRD output. Upon receiving the ground command via UHF radio, the CRD output changes the polarity of the differential voltage on the wires from -28 Vdc to + 28 Vdc. The on-state has to be continuous and above 11.5 volts for 2.5 ms minimum in order to effect the FU firing. If the signal is interrupted by noises (as short as microseconds), the time period resets and the count begins again until a full unintermpted 2.5 ms duration is met.

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Electrical Power Efficiency

As shown in Figure 2, two squib actuated silver-zinc batteries (28 V) are used to power up the Flight Safety Sys- tem (FSS) part of IFSS (including the command receiver decoder, acceleration inhibit circuit, and all FUs). This centralized battery design reduces system complexity and saves weight. The other power conservation implementation is to continuously power up FU circuits at a steady state current of a b u t 100 mA per each FU, i.e., operating in a “trigger-ready mode.” The FTOS is always armed during the missile prelaunch phase. The arming time is approximate 9 sec. Without the support of a local “on stage” battery, the FU was designed to maintain a “keep alive” status for 75 ms after the cutoff of the input power. This time allows the FU to function in a PSS scenario, because the built-in time delay from receiving PSS signal to the destruct actuation is only 2.5 ms.

FU Location

The two FUs for each solid stage are mounted on the forward skirt Y-joint region in opposite azimuth positions. Therefore, they are well protected from motor exhaust, heat irradiation from nozzle exit cones, and possible exit cone failure modes. The FU location was selected to avoid the exposure of the redundant FUs to the same local flight environments and therefore the same potential failure modes. The FUs for PBV FLSA are located near the fuel tank forward of the axial engine.

ID. Svstem Reauirements

This FTOS was developed in the early 1980’s; therefore it did not benefit from the latest versions of current mainstream specifications, e.&, MIL-STD-I 576, DoD-E- 83578A, and AFR-127-1 (References 1 through 3). It followed MIL-STD-1512 (Reference 4), which was the dominant governmental specification for electroexplosive subsystem design and testing, and SAMTECM 127- 1 (an earlier version of Reference 3), the primary requirements document for range safety hardware. They were basically similar to the specifications in References l through 3; but substantial differences do exist, as described below:

The following test requirements which are unique to MIL- STD-1512 were adopted by FTOS:

* Temperature and humidity test. Eight full cycles con- sisting of 11 hr at 126”F, 95% RH, 5 hr at -5SoF, and 3 hr wansition time - Transportation vibration (Figure 3a)

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* Temperature - altitude test Thermal time constant test for EBWI

The following MK-STD-1512 design guidelines were executed for FTOS:

- Firing circuit shall have its return grounded only at the power source. Dual-pin EBWI design was chosen over coaxial EBWI design.

* All firing circuits shall be physically isolated from any other circuit.

* Included monitor circuits. * Implemented a redundant AID switch for positive in-

terruption of the High Voltage Enable (HVE) and DC Logic Enable (DCLE) for all FUs.

The following test requirements are different from those specified in References 1 through 3 because MIL-STD- 1512 does not have these provisions:

Flight vibration and flight shock (Figures 3b and 3c): These test levels were established based on Minuteman Missile flight test data with compensation for the design differences between Peacekeeper and Minuteman and an added margin of 3.5 dB. References 1 through 3 call for a 6.0 dB margin. Peacekeeper flight data later verified that these test levels provided more than 6.0 dB margin.

* Function test at temperature extremes plus a margin of 10°C: The EBWI was tested at 174"F, while this type of test was not executed for FTOS components per

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Figure 3b. IWOS Ultimate Flight Vibration Spectrum

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Figure 3c. FI'OS Ultimate Flight Shock Sprectrurn

References 1 through 3. The Peacekeeper is launched from a silo with temperature controlled at 70 f 5°F. Therefore, the operation temperature specification is 40 to 100°F for all missile components. This small temperature range is also defined as the ambient temperature for ITOS components. References 1 through 3 specify: "The worst case temperature limits adding! subtracting 10°C margin for the test or be tested at -65 to 160'F whichever condition is more stringent."

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The choice of FTOS test temperature is very justifiable. First, the detonation explosive train of FTOS has been established to be not sensitive to such a small temperature difference. Second, heavy thermal insulation was designed and implemented for FTOS. Therefore, for the limited flight duration in which FTOS is critical, the temperature of FTOS is not much different from that in the silo.

- Electrostatic discharge (ESD) test: Elecmcal energy of 25 kV stored in a 500 pf capacitor was discharged into the EBWI in both pin-to-case and pin-to-pin modes via a 5 kohm serial resistor as prescribed in MIL-STD- 1512. However, in the later revisions of References 1 through 3, the resistor has been disallowed for the pin- to-case mode testing. Missing this new test requirement has minimum impact on the FTOS EBWI for two reasons. Fist, the exposure of the EBWI pins to human handling is minimal. The EBWI has unusually deep recessed pins, 0.375 in. from the orifice of the connector body (0 = 0.59 in.). This geometry prevents the physical contact between the pins and a human fin- ger. Second, EBWIs and FUs are shipped and stored in separate packages. This was dictated by the need for preflight inspection of these two components. After satisfactory inspection, they are mated followed by x- ray inspection for proper connection. From this point on they are handled as one unit for installation on the missile so that ESD exposure on the EBWI is elimi- nated. The EBWI was designed with a built-in short internal discharge path between the pins and the steel body away from the explosive. Therefore, the EBWI will likely survive the pin-tcr-case discharge without damage of explosive and bridgewire, if the above described test is performed.

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* EBWI spark gap breakdown test and 500V no-lire test: These tests were never performed on the FTOS EBWI because the EBWI does not have a built-in spark gap. References 1 through 3 apparently havegeneralized the requirements for one type of EBW ordnance system to all EBW ordnance systems. A spark gap is required for EBW system operation because it creates the high voltage discharge. The gap can be located either in the firing unit or in the EBWI. For the latter case, the breakdown voltage and no-fire have to be thoroughly tested as prescribed in References 1 through 3. The FTOS belongs to the former case. There is a spark gap and a triggered spark gap (Sprytron) in the FTOS FU. Their reliabilities and performance are thoroughly tested at the component level and in the FU tests. In other words, the built-in spark gap in an EBW and its associated testing are not necessary for the Peacekeeper FTOS EBWI.

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Discharge voltage and energy margin requirement: The margin is specified to be 2.0 in References 1 through 3. This requirement was met by sound engineering in the FTOS design. The minimum voltage on the FTOS FU's firing capacitor is 2400 V. The maximum all-fire test voltageofFTOSEBWIis-1200V. Thetheoretical all- fire voltage of FTOS EBWI based on Bruceton test results is -800V.

The following test requirements are unique to the Peace- keeper FTOS:

Aerodynamic heating: FTOS components are subjected to the temperature-time profile described in Figure 3d to simulate the worst case missile aerodynamic heating effect.

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Figure 3d. Heating Profile for FTOS Components

System function test after 4 hr, 0.16 psi vacuum soak: The objective was to test FU seal effectiveness

System function test at 0.16 psi vacuum post 4 cycles of pressure cycling (ambient to 0.16 psi, 20 min. dwell per each pressure level): The objective was to test FUI EBWI interface reliability at a vacuum simulating 100,ooO feet altitude.

Ultimate shock due to missile breakup (Figure 3e): This test was performed on FUs on a best effort basis due to the difficulties involved in generating and recording the high level shock.

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Figure 3e. Ultimate Shock Spectrum Due to Missile Breakup

The following test requirements are in common with those specified in References 1 through 3.

Hermetical seal: The maximum allowable leak rate is 1x10-6 standard cc helium per second at apressure differential of 7.8 psi.

* Static acceleration: 11.0 g

- Storage temperature: -5VF to 126°F

- Radiograph, electrical bonding, visual inspection, cor- rosive atmosphere, RF susceptibility, electromagnetic interference (EM).

In summary, Peacekeeper FTOS test requirements are compatible with the current mainstream ordnance specifications, References 1 through 3, and in many cases, the tests exceed that required by these references.

N. Ordnance Components Design

The EBWIs, OTAs, and TBIs used for FTOS are identical to those usedin Peacekeeper Ordnancelnitiation Sys- tem (01s). The 01s initiates nominal missile functions, e.g., stage ignitions, stage separations, Thrust Vector Con- trol gas generator initiations, extension of nozzle exit cones, and liquid and gas valve actuations. The numbers of these components in the 01s are about tenfold higher as compared to their FTOS counterparts. This implementation achieved cost effectiveness and a broadened testing database for FTOS reliability.

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EBWI (Figure 4)

Pentaerythritol-tetranivate (PETN) explosive is used both for the primary charge and output charge (density 1 .O and I .7 gkm3 respectively). The resistance of the 0.0015 in. x 0.040 in. NEYOR-G bridgewire is 0.22 f 0.03 ohm. The KOV4R pins are insulated by the alumina ceramic header. The metal-ceramic interfaces are bonded by silver brazing and silver alloy solder. The objective of the bridgewire pads is to reduce the effective bridgewire length while maintaining a proper pin-to-pin spacing for high voltage insulation. The tapered ceramic sleeves around the pins are designed for an environmental seal when they are mated with the rubber stem in the High Voltage Con- nector (HVC) on the firing unit. For interfacing with an OTA and FU, the EBWI is contained in an EBWI adapter which has threads for mating. The functional requirements are: 1) all-fwe by a high voltage discharge pulse (as defined by a 0.2 ohm resistor) of 1200A peak current and a rise time of 0.35 H.15 p.s and 2) produce a dent of 0.01 5 in. minimum depth in a steel witness plate having a Rockwell B hardness of 84 to 90.

Figure 4. EBWI and RI Contigumtion

OTA (Fieures 5 and 6)

The OTA is constructed from Confined Detonating Cord (CDC) and equipped with end tips for interfacing with EBWI, TBI, and LSC. The core loading of aluminum sheath formed mild detonating fuse WDF) in the CDC is 2.5 grains/ft of Hexanitrostibene (HNS), a high temperature resistant, insensitive high explosive. The polyethyl-

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Figure 5. CDC Construction

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Figure 6. OTA End Tip Construction

ene rubes, fiber glass, and stainless steel braids of the CDC effectively confine all explosion products when the HNS detonates. The end tips use a transition charge and booster charge explosive train also made from HNS. The end tip for EBWI and TBI interfaces has a diameter of 0.153 in. The end tip for the LSC interface has a larger end tip (0 = 0.190 in.) and more HNS. Minimum required dent depths in a steel witness plate is 0.010 in. and 0.015 in. respectively. The detonation propagation speed in CDC exceeds 6.2 km/s. OTAs for FTOS are protected by a 0.055-inch- thick RTV sleeve as a precautionary measure for thermal protection as the HNS has a temperature rating well above 300T .

TBI (Figure 7)

The TBI is initiated by an OTA end-tip coupled by a ra- chet nut. Both the donor and the acceptor charges across

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Figure 7. TBI Schematic Diagram

the bulkhead are made of high density PETN. The CuO/ Ti output pyrotechnic has a high heat content which facilitates the initiation of igniter materials, e.g., the B/KNO, pellets. For gas pressure generation, an interface booster charge of high gaseous products is required. The body of the TBI is made from a vacuum double melted stainless steel for the maximum strength and resistance to shock damage. This enables a minimum bulkhead thickness for reliable detonation transfer and strength for post fire pressure confinement.

Requirements for the TBI (when fwed into a 20 cc pressure bomb): 1) the pressure shall reach 180 psi within 3.5 ms, 2) the pressure shall increase from 50 to 180 psi in 0.75 ms, and 3) the peak pressure shall be 450 ? IS0 psig. TheTBI shall withstand an internal pressure of 8400 * 400 psi across the bulkhead for a minimum of 3 minutes at a temperature of 300 +_ 30°F with zero leakage.

LSCA (Figures 8 and 91

There are two LSCAs per stage, one installed on each side of the raceway for Stage I and Stage II. Stage III uses two half-circle LSCAs mounted near the outer edge of the forward motor dome near they-joint. The LSCAs consists of a charge holder, a retainer, an initiation housing, attachment studs. and LSC with an OTA attached.

The charge holder is made of extruded silicone rubber and is bonded to the retainer at the time of assembly. The Stages I and II retainer material is a straight aluminum channel. The Stage IU retainer is angle aluminum which has been roll formed to the appropriate diameter to facilitate the installation on the dome locations. With the holder bonded to the retainer, it provides the required stiffness to protect the LSC during handling, installation, shock, and vibration environments. It also provides an excellent thermal insulation for the LSC.

Cadmium plated steel alloy studs are pressed into the side or top of the retainer for mounting the LSCA onto the brackets in the raceway and the Stage ID forward dome.

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Figure 8. Stages I and II LSCA Schematic and Installation

Figure 9. Stage ID LSCA Schematic and Installation

In order to offset the dome expansion resulting from motor pressurization, Teflon tape is attached to the bottom of the Stage III LSCA holder to provide a low friction contact between the LSCA and the dome surface. This LSCA is rigidly mounted only at the initiation end. The rest of the brackets are designed to allow the LSCA to move when the dome expands. A cap is welded to the end of each LSCA retainer to protect the LSC end and allow connection of a grounding strap. Both ends of the LSC are welded with an aluminum foil closure disc for hermetical seal.

The initiation housing serves several purposes. It provides an attach point for the OTA, maintains the proper interface between the OTA and LSC, and provides an attachment point to the retainer. The OTA is threaded into the initiation housing and secured with a cup lock. The threaded area of the OTA end-tip is coated with polymer sealant during assembly for protection against the mois- ture entering the LSC/OTA interface.

The LSCs are manufactured by state-of-the-art extrusion and forming of aluminum tubes containing cyclonite (RDX) explosive. The core load for the Stages I and II LSCAs is 325 graindft; the load for the Stage Ill LSCAs is 40 graindft. LSCAs for Stages I and II are of identical design except the former is 11.5 in. longer because of the longer size of Stage I. The small cross section of the 40 grainslft core load requires the initiation end of the LSC in the Stage 111 LSCA to transition from its round shape to a chevron shape. This provides a better coupling with the OTA end tip.

The functional requirements for LSCAs are simple. The minimum average penetration depth in the steel witness plate(s) shall be 0.270 in. for Stages I and II LSCAs and 0.050 in. for Stage Ill LSCA at the optimized standoff distances which are built-in and maintained by the physical configuration of the charge holders.

FLSA Figure 10)

The FLSA consists of a cutting blade welded to the bottom of a piston and apinch bar. Shear pins at each end of the pinch bar retain the bar and the piston assembly in the unactuated position through any of the mechanical environments. The pinch bars are oriented to travel in the cylindrical grooves in the internal wall of the housing. The driving force to move a pinch bar and blade is provided by propellant contained in the counterbored piston. The piston has an O-ring gas seal that contains the propellant gases produced by the combustion of a high temperature resistant propellant. Upon FLSA initiation, the piston-blade assembly first punctures the fuel line to vent

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Figure 10. FLSA Schematic Diagram

the fuel from the tank. After approximately 0.125 inch of travel (and the initial puncture), the pinch bar applies force to the tube, beginning the restriction of the fuel supply to the axial engine.

The FLSA is a redundant device with two independent mechanisms in the Same assembly. These identical mechanisms are assembled to the fuel line in an opposed configuration and bolted together, through square flanges, at the interface. When activated, the FLSA punctures and pinches the Stage IV fuel line. This line has an outside diameter of 1 .OO in., a wall thickness of 0.028 in., and is made of stainless steel. The cutting mechanisms can perform singularly or simultaneously to achieve the desired results. The FLSA uses two TBIs as part of the assembly to ignite the propellant for gas generation as described previously. The functional requirements of the FLSA are: 1) The cut shall have a drain rate of 6 ft3 water under 280- 3M) psi of pressure within 77.6 sec. 2) The gap width of the crimped tube shall be 0.015 inch maximum.

V, FU Desien (Figures 4.11, and 12)

General Design Guidelines

lntemated FUEBWI Assemblv (Fieure 4L This implementation minimizes the detrimental reduction of firing current available from the FU due to the impedance loading of an otherwise required cable. It also eliminates the concern of the EMI effects due to the use of a cable. It nevertheless reduces the assembly to a one-shot device. Because with the high shock generated by EBWI firing, the FU is considered no longer flightworthy.

Electronic Technoloev. A conservative approach was adopted for FU design to achieve the required reliability. The state-of-the-art technologies of the late 1970's were used. Discrete components were chosen over the hybrid circuit. Only a few well established integrated circuits were used, e.g.. LMlO5H regulator and LM193A comparator. The active components are transistor-transistor logic based. Carbon resistors were chosen over the metal film resistors. Multilayer printed-wiring boards, connector termination boards, and flexible circuitry for interconnection were adopted.

power Reeulation, For power conservation and reliability, the high voltage circuit and trigger capacitor related circuit uses unregulated 28 & 4 Vdc input power. There- fore FU testing has to incorporate tests at both 24 and 32 V. The logic circuitry is powered by regulated 9 V from the LMIOSH output.

Lightning Protection. Each inputJoutput wire contains Zener diodes and series resistors for lightning protection. The missile was designed to be able to withstand lightning strikes during launch. This feature has never been demonstrated during a launch because of range safety constraints.

E M Considerations. The high voltage circuit uses a dedicated printed wireboard which is separated from logical circuit boards. An EMI filter attenuates the current noise spikes from the inverter circuitry that is fed back to the 28 Vdc power line.

Single Point Failure Prevention. The logic circuitry is designed such that a single electronic component failure will not produce an inadvertent firing of the unit. For this reason the circuitry has two parallel and almost identical paths and will not produce a trigger unless proper conditions are met in both of the paths.

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Figure 11. FU Block Diagram

Figure 12. Cutaway of Firing Unit

Mechanical Construction

Housins It consists of a heat treated aluminum can, a mounting plate, a flange, four card guides, and four gus- sets. An O-ring seal is used to seal the can and the flange.

Connectors. MIL-C-38999 hermetically sealed connectors are used for command destruct (7-pin) and interface connector (41 -pin). The connector for the EBWI (HVC) was custom built due to the high voltage requirement. It uses an alumna ceramichazing technique similar to EBWI to achieve the hermetical seal. The environmental seal of the interconnection housing of HVC to EBWI is provided by a silicone rubber insert.

Shock Mounts, Four mounts suppan the FU on the base plate, which in turn is secured to brackets on the missile. The implementation was necessary to meet the shingent shock environment. Each mount has a force constant of 155 Ib/in. The natural frequency in conjunction with the Fu (-8.5 Ib) is approximately 38-46 Hz. The ultimate tensile and shear strengths for each mount are 300 lb and 120 Ib, respectively.

Insulation Boot. A cast RTV insulation boot covers the entire exterior of the FU except the input connectors. The thickness of the insulation is 0.16 to 0.19 in. The EBWl adapter and OTA end tip is also covered by a thick RTV boot. The OTA is covered by a 0.055-inch-thick RTV

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sleeve. Thus the entire FTOS, including the LSC, is ther mally well insulated.

High Voltaae Circuit. This circuit consists of the follow ing elements:

. Push-pull inverter converts the 28 Vdc to a 28 V, 25

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kH2 ac.

* X22.1 power transformer converts the voltage to 619 Vac.

* X5 voltage multiplier and rectifier assembly produces 3000 V to charge up the 0.5 pf EBWI firing capacitor.

* The capacitor is connected to the HVCEBWI via a vacuum trigger switch, or Sprytron. The switch has a holding voltage of 5000 Vdc and is designed by and manufactured under the management of the Department of Energy.

Integrated with the Sprytron and connected to its trigger input is a 800 Vdc rated spark gap.

This bigger voltage (-ZOO0 V) is supplied by discharge of a 44 Vdc, IS pf trigger capacitor into a trigger transformer.

- The discharge of the capacitor is conaolled by two SCRs connected in series, one for each redundant logic circuit path.

Loeic Circuit. This circuit consists of the following elements:

- The actuation of each SCR for discharging of the trigger capacitor is controlled by the output of an LM193A comparator.

The SCR trigger from the LM193A output requires an input exceeding 3 V on the comparator which is connected to the PSS signal input and the output of the detector for the CD signal. Therefore, both PSS and CD can lire the SCR.

- Another LM193A output is connected to this PSS input and the output of theA/PSI amplifier. When t h e N PSI is in effect, this LM193A acts as a switch to short the PSS induced voltage to ground, thereby locking out the PSS signal so that no triggedtiring occurs.

- The 2.5 ms time delay for the destruct, either due to the PSS or the CD, is controlled by a RC network at the ~2

input of the comparator which is connected to the PSS signal.

* The 75 ms “keep alive” of the logic circuit is provided by a capacitor bank of 4 @ 39 pf at the input of the DCLE line.

* An intelligent “auto bleed down” circuit was implemented in the FU to dump the trigger capacitor during normal power shutdown. This is necessary to prevent the potential erratic response of the circuits from acci- dental firing when the regulated 9 V power provided by the LM105H begins to lose its effectiveness.

Monitor Circuits

The following parameters were determined to be critical for the FU operations:

The voltage on the 0.5 pf firing capacitor (HVM) for EBWI initiation: It is provided by a 1oOO:l resistive divider. Therefore the limits of this measurement are 2.3 to 3.3 V (2300 to 3OO0Volts) corresponding to the HVE input of 24 to 32 V, respectively. The divider is also acting as a bleeder for the high voltage.

- The voltage on the 15 pf trigger capacitor (TM) for Sprytron actuation: It is provided by a IOOO: 1 resistive divider. Therefore the limits of this measurement are 42 to 46 mV (4246 V) corresponding to the DCLE and HVE of 24 to 32V, respectively.

* The inhibit voltage (&I) on the input of the LM193A comparator for the A/PSI: It is provided by a SO0:l resistive divider. Therefore the limits of the measurement are 18 and 36 mV for one and two inhibit signals, respectively.

These three monitor measurements are critical. They are continuously displayed on the video screen during the prelaunch and flight test operations. Range Safety will not allow a launch unless all eight FUs are correct.

FU Signal Conditioner W S C )

The IFSS was drastically streamlined for the OT&E flights as compared to the DT&E flights. The FUSC was designed to meet the limited downstage wire pairs available and the simultaneous real time display of all three FU monitors from each of the eight FUs. It receives the monitor signals and amplifies them from the 50 mV range to the 5 V range. The multiplexing is carried out by a 80 Hz clock built in the unit. Its output switches between the

1 1 American Institute of Aeronautics and Astronautics

three monitor signals and a full range reference signal in 12.5 ms intervals.

The circuit is contained in a cylindrical housing which is strapped on the base plate of the FU. A single 22-pin connector at the end facilitates all inputJoutput channel connections, including the 2PV input power which is then regulated by a LMI 05A. The telemetry signal is decoded at Vandenberg AFB for visual display. A formal qualification was successfully performed for FUSC in 1987.

VI. Pressure Switch (Firmre 13)

The PS is an interface device between the 01s and FTOS. Its purpose is to provide apositive control on the APSI signal to the FU to prevent a PSS termination as a result of a normal missile staging operation rather thaii B true PSS. The PS was custom developed for FTOS

u a RI --I I

.I--

Figure 13. Pressure Switch Construction and Operation

~~

In ,.-An ulvrr to sense the stage separation ordnance response lo a guidance a d control command, the PS is inserted in the OTA for the staging ordn~ance initiation via an OTA “tee.” The detonation gas pressure of the OTA, attenu- ated by a small orifice in the PS, drives a pis:on in the PS, which in turn completes a dual circuit for APSI signz!~ as described previously in Section n.

The requirements for the PS are: 1) switching pressure shall be 75 to 15Opsig. 2) the switching time shall be less than 5 ms, 3) contact resistance shall be 0.1 ohm maximum, 4) the insulation resistance shs!! be 100 Mohms

minimum, 5) manual resettable post pneumatic actuation, ~~~

and 6) diode verification. L Y

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Develoument Tests

The tests were focused on the LSC because it had to bca custom fit for Peacekeeper. TBI and OTA designs were adopted from the Navy programs; therefore, no develop- ments -ere required. These activities were quite system- atic.

LSC sizing and optimization cutting tests were performed on integrated subscale motor caselinsulatorlificri propellant samples. It was concluded that a 325 graindft AV RDX LSC could sever a Kevlar case up to 1.2 in. thick at 2 standoff height of 0.470 in. Similarly, a 40 grainslft AV RDX LSC could sever a Kevlar case up to 0.250 in. thick at an optimized &off height of 0.230 in. These LSCs were chosen as the baseline de&gn. The equivalent penetration in steel witness plate was detemixd to be 0.270 in. and 0.050 in., respectively.

Propellant safety tests used samples and a setup similar LV he above test except that real stage propellants were used. The results indicated that breakup and partial initiation occurred for the prope~~ants of all three stages, but no detonation was observed.

Full scale motor destruct tests were performed for all !Ire solid stages at Phillips Laboratory at Edwards AFB. Base- line LSCAs were used to cut each stage at about 2.5 sec into the motor firing, a time which offered the worst case scenario because both propellant mass and chamber pressure were maximized. The test of Stages I and EI pro- ceeded as planned, the jet from t!!e LSCAs severed the case and the chamber pressure extinguisfid in millisec- onds. Propellant breakup and partial combustion were observed, but the explosive yields were very low, of the order of IO and 55 Ib for Stages III and I, respectively, indicating that detonation did not occur. The Stage Il testing was a no-test&ecause the motor was defective. It burst before the planned LSCA actuation. (Reference 5).

In the early 1980’s. EBWIs, OTAs, andTBIs were manufactured in support of stage testing, OIS, and FTOS flight proof tests (FPT). Lot acceptance tests(LATs) were performed on these components and the FFTs were executed. Several prob!ems were encountered, and after extensive investigations, corrective^ actions were implemented as necessary. ~ ~~

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12 American lnstiture of Aeronautics and Astronautics

Blow out of OTAs near the end tips were observed in several OTAs in 01s F'PTs. It was concluded that it was caused by the insufficient free volume in the ordnance connections, is . , TBIs and CKA "tees." No corrective action was implemented as no detrimental effect of blow- out was identifiable.

Failure to propagate occurred in one OTA near one end tip in an 01s F'PT. (This was one of only three OTA function failures Peacekeeper had ever encountered out of approximately 4600 CWA firings.) Investigation revealed that the IS-hour combined transportation and preflight vibration test time had caused a fatigue fracture of the aluminum sheath around the HNS case. As the corrective action, the test duration was reduced, since it was a gross overtest. The transportation vibration test was modified to allow the use of packaged OTAs.

One EBWI failed to initiate in the LAT. (This was the only EBWI function failure Peacekeeper had ever encountered out of approximately 3300 EBW firings). Investi- gation revealed two possible causes: 1) the gap between the initiation and output charges was too large and vari- able, and 2) a vacuum condition was induced in the EBWI by the newly implemented vacuum welding process. Both problems were resolved by tightening the assembly con- trols and changing the welding process.

v

01s and FI'OS Flieht Proof Tests

Besides the OTA failure described above, 27 TBIs, 20 EBWIs, and 41 OTAs were later successfully tested in 01.7 FPT. Ten Stage I LSCAs, 10 Stage III LSCAs, 10 PSs, 19 FLSAs (including redundant TBIs and OTAs) and 3 FUs were used in FTOS FF'Ts (1981) in which they were subjected to the environments in Section El. In addition, one FU was subjected to the EMI testing per MIL- S T I 4 6 1 and 4 6 2 (References 6 and 7). Two problems were identified and corrective actions were implemented:

- Many EMI failures occurred. These led to the redesign of the EM1 filters in the FU.

* During shock and vibration, the PS exhibited chatter- ing in electrical contacts. This led to the redesign of the PS. Successful make-up testing was completed.

The overall testing was successful. A couple of apparently serious failures were diagnosed as test errors. In F'PT the components were tested as assemblies. The firing tests were largely initiated by test OTAs, which are W A S with either output end tip type initiated by an electrical detonator instead of EBWI. \d

Qualification Tests

The FTOS system qualification was performed in 1982. Thirty-three Stage IAl LSCAs, 33 Stage 111 LSCAs, 33 FLSAs, and 32 FUs were tested either in assemblies initiated by test OTAs or integrated with FUs. The Stage Yn LSCAs were subscale units, i.e., the same end designs as the full scale units but with shorter length. This was necessary to accommodate the test equipment. The environment requirements in Section III were used. All tests were successful except 2 OTA lines showing propagation stop- page. (These were two of the only three OTA function failures ever encountered in the Peacekeeper program out of approximately 4600 OTA firings.) The cause and dis- position were the same as in the OTA development test described above. EM1 tests were performed on three FUs. They passed MIL-STD461and 4 6 2 requirements, except there was a minor excursion in two frequencies in CEO3 and CEO4 tests of the FU power return line. The test series was considered also as qualification for the LSCA, FLSA, and FU.

The EBWI has gone through three qualification tests:

In the 1982 qualification, 180 units were tested; 130 units were environmentally tested and 50 units were successfully tested for the RF environments of 1 -W, 5- sec exposures at nine frequencies. Post the environmental exposure, about 40% of the units failed the hermetical seal test, but all units successfully functioned.

In 1985, the EBWI with improved brazing procedures was subjected to a requalification. Two hundred units were successfully tested, including the post-environment hermetical seal test.

- In 1986, 200 E B W s from an alternate source were tested for source qualification. The test was successful except twounits failed the post-environmental hermetic seal test. A decision was made to implement a 40 cycle temperature cycling (45'F to 126°F) test both for header subassembly stress screening and post-manu- facture screening. This implementation has proven to be effective.

The TBI has gone through two qualification tests.

In the 1985 TBI alternate source qualification (100 units), two units exhibited a slightly longer time delay to 180 psi, 4.5 ms versus the 3.5 ms maximum time allowed. An investigation concluded that an excessive air gap between the acceptor PETN and the CuOlIi output mix was the cause of the anomaly. As correc-


live actions. the loading pressure of CuOmi was in- creased and the excessive free volume was reduced. The modified design later passed a repeat qualification.

The LAT for explosive components and PSs includes the nondestructive inspection on dimensions, workmanship, x-ray and n-ray inspections, and hemetic seal test. For the EBWI, bridgewire resistance, insulation resistance, and thermal transient properties are also measured. These activities are applied 100% to the lot. Destructive tests at ambient temperatures of 62 to 95°F are conducted on 6% or 30 units minimum sample size of each EBWI lot and 10% of each TBI, OTA, LSCA, and FLSA lot. The functional requirements are shown in Section N.

Pneumatic actuation was adopted as a nondestructive test at the factory for the PS at a 10% lot sample size.

Certain testing has to be performed during the manufacturing of the components. Examples are the hermetic seal test for the LSC closure discs and the OTA end tip booster charge cup which were welded before the final assembly. The FU in-process test consisted of the following four tests which were performed on all FUs:

Non-operating temperature cycles: -63'F to I3 I "F for a minimum of 2.5 hr of dwelling and a temperature change rate of 6 f 3°F under a power-down condition.

Operating temperature cycles: 3 cycles from 32 to 105°F for a minimum of two hr dwelling and a temperature change rate of 6 + 3°F under a power-up condition.

Vibration test: Two minutes of 8.9 grms of vibration as a workmanship verification test.

* Low pressure test: The FU shall'be subjected to a pressure of 0.03 psia for 30 min under the power-up condition and then returned to ambient pressure. This test was implemented to replace the originally required helium leak test on the FU seal. The latter test was diff- cult to perform successfully because there are numer- ous micro voids in the O-ring groove and the connectors which trap helium gas and appear in the test results as a false large leak rate.

The acceptance test procedure (ATP) for the FU is performed at ambient temperature only. The test procedure consists of the following elements:

* Independent variables: - HVE voltage: 24 and 32 V - DCLE voltage: 24 and 32 V - CD voltages: 10.0, 11.5.24 and 32 V - Presence of one or two O S 1 voltages - PSS resistance at 50 ohms and 25 kohms

* Constant inputs: O S 1 (l), O S 1 (2)

* Dependent variables: - Current drawn due to the application of HVE and

- Monitor voltage readings: HVM, TM, and IM - Destructive Output Pulse (DOP) characteristics as

measured by a FU Output Monitor (FUOM), which is a resistance divider with total resistance of 0.2 ohm and ratio of 100:l. The characteristics are the peak current, rise time, and pulse duration.

DCLE, i.e., IHVE and IDCLE

* Static measurements - Input resistance between pins in the connector - Reverse polarity current for HVE, DCLE, CD, APSI

( I ) and A/PSI (2).

* Dynamic measurements - Input signal current for PSS, CD, APSI (1) and A/

- DOP delay time - Keep-alive time - HVM decay time after HVE was turned off - Trigger capacitor charging time

PSI (2).

* Function verification - Command destruct, DOP issued - Premature separation, DOP issued - PSS Inhibit, DOP not issued

The measured performance is accepted or rejected based on limits established from the FU test database gathered from testing hundreds of FUs.

In addition to the LATs of the components, a LAT of FTOS at the system level was implemented. The FTOS was delivered on a missile kit basis which included eight FUs, two LSCAs for each stage, an FLSA, pressure switches, and miscellaneous components such as nuts and adhe- sives. A lot of mOS kits was defined as a maximum of 15 m0S kits. A kit LAT included one half of a redundant FTOS for each stage (one FU and one LSCA for each stage), a full FLSA (two OTAs and two TBIs and FLSA) and 10% of each PS production lot. All of the items were nondestructively inspected and destructively tested at ambient conditions.

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14 American lrrstitute of Aeronautics and Astronautics

M. Aging and Surveillance Tests

The AIS program for FTOS was formulated by the following considerations:

* The required FTOS service life was projected as IO

d

years.

The desired service life for Peacekeeper was 25 years

The production FTOS hardware was delivered in a time span of approximately 7 years starting in 1986.

Therefore, the original 25-year A / S program adopted the following strategy:

Stock a large quantity of FTOS components in the initial delivery for A / S .

Perform nondestructive tests every year on these com ponents.

Perform destructive functional tests on the component samples at intervals and increase the frequency as time goes on.

From the test results, perform analyses to determine the aging trend if any.

W'

* predict the performance of younger components from analytic results on these oldest components.

* Use the oldest FTOS hardware fust in tests to warrant the performance of the younger hardware.

In time, test younger hardware to provide information on lot-to-lot variations.

Currently, we are eight years into the program of testing the FTOS ordnance components manufactured in 1986; no aging trend has been detected so far. W~th the new scope of the Peacekeeper program, which will terminate flight tests after 1996, we are confident that FTOS will reliably support all OT&E flight tests in the plan. If the Peacekeeper flight test program is extended, we can con- tinue the AIS program for years to come in accordance with the original plan.

There are two potential age-sensitive Components in the FU which need AIS observations, the spark gap and the Sprytron. The former is filled with sub-atmospheric pressure inert gas; the latter is a vacuum tube. Gradual leakage of air into these devices may drastically affect their characteristics. The original mOS AIS plan required

yearly testing of 1 2 N S FUs. In 1993, due to the fact that range safety requires the repeat of FU ATP at Vandenberg AFB within the three months prior to the flight test and that no FU aging trend was notable in seven years of testing, a decision was made to terminate AIS testing of the FUs. The comprehensive preflight ATP test data and the flight data provide a meaningful indication of the A/S of the FU design.

X. ITOS Performance

Eachof 18 DT&EIOT&Edevelopment flights (FlightTest Series, FTS) and 18 FOT&E flights (Glory Trip, GT) carried an FTOS. Due to the success of the Peacekeeper test program, missile destruct was executed only once. In August 1989, during the early period of Stage III operation, a gyrocompass assembly anomaly led to the command destruct of the missile. The LSCAs on Stage III successfully terminated thrust of the stage and the FLSA was successful in disposing of the fuel in the PBV.

In early flight tests, the FLSA was actuated routinely to assure the demise of PBV propulsion capability. Recently, it has become an optional event depending on the level of excess fuel after the PBV maneuvers. The FLSA will not be actuated if the fuel reserve in the fuel tank is near ex- haustion. In any event it is the most frequently functioned component in the FTOS during flight tests.

The entire FTOS is installed on the missile at Vandenberg AFB, whereall Peacekeeperflighttestsarelaunched. ATP is repeated there for each FU prior to the flight test, usu- ally within the required three-month time period. The test is performed by using FUTS and FUOM. Eleven units are tested for each flight, eight for installation and three for backup. Over the years, over4OOFUs have been tested. Only three FUs were rejected due to precautionary measures. They had minor problems. In one case, an input filter capacitor was shorted to the FU case. In another case, a diode was wired with the m n g polarity. None of these problems would have caused a FU failure to oper- ate. This indeed is a remarkable record, since the test is very thorough and the FUTS is designed with the guidelines that it can fail a good FU but can never pass a bad FU.

The pressure switch is a component which can be nondestructively actuated. The test is not conducted at Vandenberg AFB because the PS is a highly reliable device.

The FUSC is test-verified in the stage level tests after the FTOS has been installed. The proper function of the FUSC is indicated by the reading of the three FU monitor chan-

15 American Institute of Aeronautics and Asrronaufics

nels after the FUs are powered up by the Stage Test Set. An Electrical Continuity Test Set (ECTS) which checks out the pin-to-pin resistance for the entire missile cable system provides continuity measurements for the RISC. The GT prelaunch testing showed some variability in ECTS reading for FUTS that the critria for the limits of the reading had to be established based on accumulated data.

Perhaps the single most serious incident involving an ITOS occurred during ITS-16 prelaunch activities in November 1986. In the PBV stand-alone test after the FUs were installed and H E was active, an EBWI was inadvertently initiated at about 2.5 min. into the power- up test. Because the DCLE was not activated at the time and the EM1 stimuli in the test facility were proven later to be minimal, this incident pointed to a premature breakdown between the plate and the cathode of the Sprytron which holds the high voltage across the EBWI firing capacitor. An internal breakdown in the Sprytron was ex- tremely remote as this type of device had a very successful record at DOE in which no failure was encountered in six million tests. Breakdown could also occur if the interface between the ceramic body of the Sprytron and the conformal coating was contaminated or a partial vacuum was created by stress or thermal stress. (A foamed conformal coating was used to embed the entire electronics inside the FU.) Retest of this problem FU at the factory could not duplicate the failure mode nor could a com- plete disassembly and falure analysis of suspect components identify the cause. Thus the investigation was not conclusive.

As a corrective action, a high voltage screening test was instituted at Vandenberg AFB. The test consists of artifi- cially increasing the voltage across the Sprytron by 1M)o V above that normally being provided (i.e., at 3500 V) and leaving the tube under this voltage stress for a period of half an hour. Both the voltage and the duration of bias would accelerate any latent cause for the breakdown. Since this implementation, over 200 FUs have been successfully tested. This failure mode has never been encountered since FTS-16, neitherduring the lOOOV screening test nor during the normal FW power-up operation, indicating that the incident was indeed a very rare occur- rence.

XI. FUTS AccomDlishments

The efforts in FUTS design and evaluation deserve a special reporting because they reflect the trends of testing high reliability equipment in general and flight safety systems in particular. During the span of 18 years for FTOS development, manufacturing, and utilization, electronic

w

technology has made rapid progress. Reflecting this progress, three generations of FUTS were designed, manufactured and used. They are briefly described below.

Test Interface Unit (TIU)

The testing of the FU described in SectionVIII is simple in principle but complex due to the multiple tests required by the combination of many input variables and parameters. A dedicated test set was required. A TIU was fab- ricated to support early FU testing. Due to its portability, it continued in use for field tests and testing of F U s in environmental laboratories long after more advanced FUTSs were available.

The design of the TIU was simple. It basically consisted of power supplies of fixed voltages. The voltages were connected to the FU by wireskables and controlled by switches for HVE, DCLE, CD, AmSI ( I ) , ARSI (2). PSS (l) , and PSS (2). The value of HVM, IM, and TM were displayed on threedigital voltmeters. Thus, both test command and monitor reading recording were manually performed. This primitive design was sufficient to verify FU operation at a nominal input voltage of 28 V, e.g., the premature separation firing, command destruct firing, and inhibit no fue. The tiring can be indicated by both the HVM and TM readings. The DOP pulse shape could be monitored by recording the output of the FUOM on an oscilloscope.

Old FUTS

In 1981, a computer based FUTS was designed and fabri- cated for the in-depth acceptance testing of production FUs and the preflight FU validation at Vandenberg AFB. Two systems were built.

The computer establishes FU input conditions via several hundred relays, performs voltage measurements (the currents are measured by current viewing resistors (CVR) built in the system), digitizes and analyzes the data, com- pares it with the limit values stored in the memory, and determines the pasdfail of each test in an automated mode. Because of the short duration of the DOP (- I w), a transient digital recorder is used to store the data for analysis later on. Programmable and fixed output power supplies are built in.

Hardware and software of late 1970s vintage were used for this FUTS, including an HP Series 200 Model 35 computer (HP9835) system including keyboard, dual disc drives and printer, and HP6940A multiprogrammer. The latter data acquisition, control, and test system consisted of two 4K memory banks, two analog to digital convert-

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American lnsritute ofAeronautics and Astronautics

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ers, a timer, a counter, a monitor board, power supplies and two resistance boards for interfacing with two programmable power supplies. The relays were commercial grade. Wire wrap boards were used to facilitate changes. A FU simulator with fixed input/out impedance and voltages was built for performing the ATP on the FUTS. At Vandenberg AFB, several test-only FUs were always available for the purpose of FUTS validation. The test results obtained from these units were always consistent. The FUTS also has a self-test program built in, the Equip- ment Verification Procedure. These features helped stan- dardize the maintenance of the FUTS.

The FUTS testing sequence flow was as follows:

* Static tests

W’

Resistance measurements Reverse polarity measurements

Dynamictests Performance characteristics @24 Vdc Performance characteristics @32 Vdc Command destruct output @24 Vdc Command destruct output @32 Vdc Premature stage separation @24 Vdc Premature stage separation @32 Vdc Minimum voltage - no tire Input signal current @ 24Vdc Input signal current @32 Vdc Command destruct no fire @ 24 Vdc Command destruct no tire @ 32 Vdc

d

IO00 V high voltage screening test

The dynamic portion of the test involved 25 specific tests in a two-hour test duration. The test duration includes a total FU power-up time of about 10 min with eight tir- ings.

The data output of the FUTS was available in three for- mats: - Printout of pass or fail for each test.

A summary sheet of measured parameters along with the maximum and minimum limits.

* The detailed data stored in the computer discs can either be printed or further processed, e&, to determine the average values of the measurements and identify their maximum and minimum values.

The detailed test conditions and results are:

- Continuity test: resistance of lopairs of key pin-to-pin ‘,4’ measurements.

Reverse current test: currents for HVE, DCLE, C D , N PSI (1). and AmSI (2).

- High voltage enable test: IHVE, HVM decay time, HVM charging time, and HVM at voltage of 24 V and 32 V with DCLE off.

* DCLE test: IDCLE, TM, and keep-alive time at voltage of 24 V and 32 V with HVE off.

Input signal current test: Current of PSS ( I ) , PSS (2). CD, APSI (I) , and APSI (2); resistance of PSS measured at AmSI voltage of 5.0 V with CD voltage at 32 V and DCLE voltages of 24 V and 32 V.

Inhibit monitor test: Voltages of AmSI ( I ) , APSI (2), and IM at DCLE voltages of 24 V and 32 V.

Trigger monitor test: TM charge time, TM voltage, and HVM voltage at HVE voltage of 32 V and DCLE voltages of 0 V and 32 V.

* Destruct output pulse, command destruct, and no-fire test: CD voltage, DOP rise time, DOP width, DOP peak current, and DOP delay time at HVE/DCLE voltages of 24 V and 32 V.

Destruct output pulse, PSS, normal stage separation, and no-fire test: DOPrise time, DOP width, and DOP peak current with PSS and PSS(2) open, AmSl (I) and AmSI (2) off, and HVElDCLE voltages of 24V and 32 V.

The two FUTS were successful in supporting the production of approximately 800 FUs at the factory and pro- cessing approximately 400 FUs at Vandenberg AFB. However, they did have some shortcomings:

Because the CVRs did not generate sufficient signal levels, data resolutions were coarse. Typically, between the maximum and minimum limits of the measurement, only 4 to 6 bits were available.

The grounding of the system was not perfect. DOP generated noise could cause some random DC shift in the DOP signal, which appeared as an apparent “under shoot.” This phenomenon caused the computer to mis- interpret the signal and repon it as a test failure. Re- tests were necessary.

The commercial grade relays and the wire wrap contacts produced intermittent failures at times.


New FUTS (Fieures 14a and bl

Because of the degradation of the old FUTS and because HP had terminated production of the HP Model 35 computer in 1984 and might cease to provide service, a decision was made in 1987 to design and fabricate a new FUTS. The following new technologies were adopted:

* HP PC-305 computer, designed specifically for instrument conkol. It had an 8088 processor, 640 Kbyte RAM, ROM-BIOS, monitor card, a 20 Mbyte hard drive, a 3.5 in. 2 Mbyte disc drive, and BASIC Language Processor Card (VIPER) which uses a 68020 microprocessor, IEEE-488 Bus interface, BASIC Boot ROM, and 4 Mbyte RAM.

w

INTrRFACE PANEL

PRINTER ,

COLORADO DATA SYSTEMS CHASSIS AND

CARDS 1

MONITOR

:OMF'UTER I

10 VOLT POWER SUPPLY

/ +60 VOLT do VOLT POWER POWER SUPPLY SUPPLY

Figurr 14a. FUTS Front View

AC POWER ' DISTRIBUnON INTERCONNECT PANEL

PANEL

Figure 14b. mTTS Rear View

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18 American lnsrirure ofAeronautics and Astronautics

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- Instrument-on-a-card technology built by Colorado Data Systems. -

* Matrix Measurement Bus providing two-wire or four- wire measurement between 96 instrument or test points. The bus is comprised of six matrix relay cards each containing 32 DPST relays.

* A comprehensive system interconnect panel

Programmable gain amplifiers and resistances

* Built-in and switchable simulated loads

Digital strip chart recorder

Laser Jet Series Il Printer

The basic test sequence is similar to the old FUTS; however, the command level table is much more complex, allowing more interactive testing between parameters. The data analysis capability is more sophisticated to deal with transient changes: e.g., the time to reach steady state is measured for each change of level and compared to the specified value. The steady state is verified in respect to the specified limits and violations are reported as failures. The new FUTS has a data handling speed tenfold faster than the old FUTS. Therefore, for the same two- hour testing period of each FU, tenfold information is obtained. Typical data is shown in Figures 15 a through d.

Because of the dramatic increases in accuracy and reso- lution, the variations of the test results are much smaller than those measured by the old FUTS. The choices of the acceptance limits had to be reevaluated and correlated to the old FUTS, because the existing FUs had all been accepted per the old FUTS. The limits for the transient and interactive measurements had to be established. These issues combined with the intrinsic system complexity, consumed considerable amount of time and resources. In 1994 after a thorough fine tuning involving hardwardsoft- ware changes, the system was qualified for testing of flight FUs. Currently, both new FUTS are operational at Vandenberg AFB for flight support.

XII. Summary

From the foregoing history of design, development, testing, and performance, one can realize that the Peacekeeper FTOS is a well established and reliable destruct system. Utilization of it in other missiles and launch vehicles can lead to good service and cost effectiveness. It is encour- aging to note that the FTOS Stage I LSCAs have been

-

..--,

Figure 15% Typical FUTS Output Record: 32 Vdc Test

I -- I - - I

I *._ ........ m "-At... "ac/DI". ....... m" .m -r mc. E ....... M W L U C

r2r.t I N - r ............ I -.e ...."_I ............. L.000 N I I mriz-1 WI ......... 12. *. IO .*,Dl"

MC,PI". WC/DI*

LO .e*

Figure 1%. Typical FUTS Output Record: HVE and HVM


=.e. A ....... _U "OLt... *." I c.72 N m n e 7r.C. 6 ....... n "0ir.Q.

,Ij.% -..-t ............ L Y.f "..-t ............. ..o*o NU " * . L . ( n U L Y L ......... ,I. s.0. IO sr,*."

Yo<:o." Aac O i l . "&</Si"

> o , 0 0 3 . e * ......

Figure 15c. Qpical FUTS Output Record: DCLE and TM

successfully flown on aTaurus Vehicle and that the Lock- heed Launch vehicle, is currently in a process to use the Peacekeeper FTOS design. U On the other hand, as has been briefly mentioned in this paper, the ever progressing technology does offer areas for further improvement of FTOS. Examples are:

- Hybrid electronics packaging is now a mature technology. Utilizing it for packaging the logic circuits can significantly reduce FTOS F'U size while increasing performance.

- The molded and machinable S-Glass has a stronger bonding strength (-200 Ksi) to metals than alumina ceramic (30 Ksi). Therefore, it is an attractive material for EBW pin insulation with improved performance in hermetical seal.

* Newly emerging flexible confined detonating cord is an improved version of the Peacekeeper OTA. Prelimi- nary information indicates that it has a better resistance to vibration damage due to its smaller size and rugge- dized consbuction.

Figure 15d. Typical FUTS Output Record: DOP v

20 American lrisrirute of Aeronaurics and Astronautics

In the Titan IV Solid Rocket Motor Upgrade program, a new TBI was successfully qualified. It uses HNS instead of PETN as donor/acceptor charges and BI KN03 pyrotechnic instead of the CuOfli mix. It also offers better temperature stability.

u

XIII. Acknowledeements

Thiokol COT. hasbeen theprimecontractorofFTOS from the beginning of the Peacekeeper program development. The components manufacturers were as follows: Fu, FIJSC, PS, and FUTS: Quantic Industries; EBW. EG&G Inc. and Hi-Shear Technologies Inc.; TBI and OTA: En- sign-Bickford Aerospacc Co. and Lockheed Missiles and Space Co.; LSCA: OEA Aerospace Inc.; and F L S A Teledyne McCormick Selph Co. Hundreds of people con- tributed to the FTOS Program. To name a few, USAF Col. (Ret.) C. M. Swager, Lt Col E. C. Y. Hu, Maj. A. Reinhardt, Maj. W. T. Greer, and Capt. K. Heranza; Thiokol: Messrs. J. R. Thurston, L. D. Berchtold, R. S. Tappan, J. E Strahm, M. E Sears, R. W. Coleman, E L. Duce, R. W. Dam, and D. M. Beebe; Western Range Safety: Messrs. D. Leistico, K. Day, and M. Gotf?aind; MMTI: Messrs. C. Getzoff, M. Blcdgett, and T. Moms; TRW Messrs. J. C. Metcalf, V. J. Menichelli, and B. A. Snyder. A special thanks is due to Mr. Duce for supply- ing important archive information for this paper and com-

2 ments on the manuscript.

XIV. References

1) Mn-STD-1576, “Electro-explosive Subsystem Safety Requirements and Test Methods for Space Systems:’ USAF, 31 July 1984.

2) DOD-E-S3578A, “Military Specification, Explosive Ordnance for Spacevehicles (Metric), General Speci- fication for,” USAF, 15 October 1987

3) “Range Safety Requirements,”AFR 1 2 7 - 1 . W 127- 1,andERR 127-1, 1995.

4) MIL-STD-1512,“Electroexplosive Subsystems,Elec- mcally Initiated, Design Requirements and Test Meth- ods,” USAF, 21 March 1972.

5) “MX Stage Destruct Test,” S. F. Bridges, AFRPL TR- 82-068, Air Force Rocket Propulsion Lab, USAF, December 1982.

6) MIL-STD-461B, “Electro-magnetic Emission and Susceptibility Requirements for the Control of Elecno- magnetic Interference,” Department of Defense, 4 August 1980.

7) MIL-STP-462, “Electro-magnetic Interference Char- acteristics, Measurement of,” Depanment of Defense, 31 July 1967.

21 American lnstitute of Aeronautics and Astronautics

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