NASA Technical Reports Server (NTRS) 19960002335: Flight demonstration of flight termination system and solid rocket motor ignition using semiconductor laser initiated ordnance

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by  NASA Technical Reports Server (NTRS)

KNASA-TH-111093) FLIGHT N96-12343
DEMONSTRATION OF FLIGHT TERMINATION
SYSTEM AND SOLID ROCKET MOTOR
IGNITION USING SEMICONDUCTOR LASER Unclas
INITIATED ORDNANCE Progress Report,
1993-1995 (NASA) 22 p
G3/20 0071550
AIAA 95-2980
FLIGHT DEMONSTRATION OF FLIGHT
TERMINATION SYSTEM AND SOLID ROCKET
MOTOR IGNITION USING SEMICONDUCTOR
LASER INITIATED ORDNANCE
N. Schulze, NASA Headquarters
Washington, DC
B. Maxfield, NASA Wallops Flight Facility
Wallops Island, VA
C. Boucher, The Ensign-Bickford Company
Simsbury, CT
Copyright © 1995 by the American Institute of Aeronautics and Astronautics, Inc. No copyright
is asserted in the United States under Title 17, U.S. Code. The U.S. Government has a royalty-
free license to exercise all rights under the copyright claimed herein for governmental purposes.
All other rights are reserved by the copyright owner.
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Joint Propulsion Conference and Exhibit
July 10-12,1995/San Diego, CA
For permission to copy or republlsh, contact the American Institute of Aeronautics and Astronautics
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ABSTRACT
Solid State Laser Initiated Ordnance (LIO) offers new
MMC The Martin Marietta Company (now
technology having potential for enhanced safety, reduced
Lockheed Martin)
costs, and improved operational efficiency. Concerns
OSC Orbital Sciences Corporation
over the absence of programmatic applications of the
OTA Ordnance Transmission Assembly
technology, which has prevented acceptance by flight
Q Vehicle dynamic pressure (pounds per
programs, should be abated since LIO has now been
square foot)
operationally implemented by the Laser Initiated Ord-
SNL Sandia National Laboratory
nance Sounding Rocket Demonstration (LOSRD) Pro-
TBI Through Bulkhead Initiators
gram. The first launch of solid state laser diode LIO at
WFF NASA Wallops Flight Facility
the NASA Wallops Flight Facility (WFF) occurred on
March 15, 1995 with all mission objectives accom-
1.0 PROGRAM DESCRIPTION
plished. This project, Phase 3 of a series of three
NASA Headquarters LIO demonstration initiatives, ac-
Background
complished its objective by the flight of a dedicated, all-
Laser Initiated Ordnance has for many years been
LIO sounding rocket mission using a two-stage Nike-
claimed to offer advantages to the launch industry and
Orion launch vehicle. LIO flight hardware, made by The
for payloads. But due to lack of operational use, it has
Ensign-Bickford Company under NASA's first Coopera-
not been employed. Thus, a flight demonstration be-
tive Agreement with Profit Making Organizations,
came important to the aerospace industry in order to
safely initiated three demanding pyrotechnic sequence
provide that key operational experience and the critical
events, namely, solid rocket motor ignition from the
flight test data to surmount a major hurdle of LIO ac-
ground and in flight, and flight termination, i.e., as a
ceptance by flight programs. One concern may be at-
Flight Termination System (FTS). A flight LIO sys-
tributed to the uncertainties associated with defining
tem was designed, built, tested, and flown to support
safety requirements, including range safety. Without
the objectives of quickly and inexpensively putting LIO
the need established by a specific mission tied into an
through ground and flight operational paces. The hard-
actual flight LIO system, the generation of safety re-
ware was fully qualified for this mission, including
quirements has been difficult.
component testing as well as a full-scale system test.
The launch accomplished all mission objectives in less As a means to counter those problems, the initial
than 11 months from proposal receipt. This paper con- thought for this demonstration program was conceived
centrates on accomplishments of the ordnance aspects of by the NASA Headquarters author. The NASA Sound-
the program and on the program's implementation and ing Rocket Program appeared ideally suited to carry out
results. such a mission. The launch vehicle had to expose a
LIO payload to a rigorous launch flight environment at
While this program does not generically qualify LIO for
very low cost. Further, program activity had to be
all applications, it demonstrated the safety, technical,
maintained at a high level of quality and technical rigor.
and operational feasibility of those two most demanding
The focus of the organization performing the program
applications, using an all solid state safe and arm sys-
had to be one having technology transition interests in
tem in critical flight applications.
order for the program to receive the required priority and
detailed attention.
ACRONOMYNS
The NASA Sounding Rocket Program at the Wallops
ARC Atlantic Research Corporation
Flight Facility had the unique capability of launching a
DoD Department of Defense
payload into the desired environment at a very low cost,
DoE Department of Energy
EBCo The Ensign-Bickford Company making it ideally suited to this mission; and, hence, it
was pursued. While the Sounding Rocket Program is
ETS Energy Transfer System
predominantly associated with serving the scientific
FCDCA Flexible Confined Detonating Cord
research community at the present time, in the early
Assembly
years of NASA, and in the NACA before then, it served
FTS Flight Termination System
mainly NASA's aeronautics programs. Further, sound-
GSFC Goddard Space Flight Center
JPL Jet Propulsion Laboratory ing rocket vehicles have been used to demonstrate and
flight qualify new technologies many times; but this
JSC Johnson Space Center
involvement in a technology transfer program expands
LDFU Laser Diode Firing Unit
LID Laser Initiated Detonators that activity in a relatively new role.
LIO Solid State Laser Initiated Ordnance The initial Headquarters suggestion was formally made
LOSRD Laser Initiated Ordnance Sounding Rocket on September 1, 1993 when discussions were held at
Demonstration Program the NASA Wallops Flight Facility to explore concept
American Institute of Aeronautics and Astronautics
feasibility. The conclusion was that the demonstration secondary objective, the program used this unique op-
project would be important in bringing forth the tech- portunity to gain data from thrust termination using an
nology and would be feasible for a sounding rocket ordnance design which ports motor chamber pressure
flight. On October 21, 1993 a follow-up meeting was from the head end of a solid rocket motor, the subject of
held at the Wallops Flight Facility to advance program a separate report.
definition. Then, it was concluded that a dedicated two-
Several mechanisms were provided to verify program
stage vehicle, the Nike-Orion, would be the best choice
objectives. LIO safety was particularly important be-
for meeting program objectives and that the availability
cause safety has been stated to be a key LIO technology
of telemetry was important. NASA would provide a
feature. Further, it is clearly one feature required to
platform(s) for mounting the required LIO hardware. On
make possible the practical implementation of LIO.
January 5, 1994 a Pre-Project Initiation Meeting was
Confirmation that the detonators have fired was to be
held at the WFF to accomplish initial planning. The
determined by visual aids, radar, or accelerometer data
conclusion was that the desired quick, low-cost LIO
telemetered back to WFF. Adequate flight data from the
motor ignition and flight termination system flight
mission was necessary to transition the results of this
demonstration goal would be feasible, and WFF com-
flight to industry.
mitted to performing the mission.
Program Implementation
Definition
The program management approach was key to our suc-
"Laser Initiated Ordnance," that is, "LIO," as defined in
cess with accomplishing the tough program objectives.
this program, uses laser light energy to replace electrical
It was decided that to be successful this program must
energy for initiation of ordnance and uses all solid state
use a highly leveraged technical and programmatic ap-
electronic components throughout, including the safe
proach if we were to effect realistic results with the
and arm system. The other aspect of laser initiated ord-
cost, performance, and schedule constraints.
nance, distinguishing it from conventional ordnance
systems in general, is that a fiber optic cable replaces Hence, relevant hardware pedigree was essential, requir-
the ordnance cord to transfer energy. ing the use of existing designs. Ordnance interfaces had
to be maintained; otherwise an unaffordable qualification
Two laser technologies have been developed, rod and
program would become necessary. A management
diode. Rod laser ordnance ignition has been previously
oversight and advisory approach was needed to provide
demonstrated on the Small ICBM Program. With the
the best possible expertise to prevent costly errors. To
rod laser ordnance initiation experience behind us, the
that end a government-industry team was organized
laser choice for this NASA flight demonstration pro-
comprising several NASA centers (Wallops Flight Fa-
gram was, thus, the laser diode. In this program we
cility, Johnson Space Center (JSC), Jet Propulsion
refer to "LIO" strictly as solid state laser diode technol-
Laboratory (JPL), Goddard Space Flight Center)
ogy, unless otherwise specially noted.
(GSFC) plus the Department of Defense (The Aerospace
Objectives Corporation), Department of Energy (Sandia National
Laboratory), range safety staff (Wallops Flight Facility,
In order for this to a be quick, results-oriented program
eastern, and western ranges), The Ensign-Bickford
and one performed at low cost, it was designed to be
Company (EBCo), The Martin Marietta Company (now
highly focused, that is, a program to resolve the key
Lockheed Martin), Orbital Sciences Corporation (OSC),
issue of demonstrating the programmatic feasibility of
and NASA/DoD/DoE Aerospace Pyrotechnics Systems
all solid state LIO.
Steering Committee. That advisory group was referred
A high risk, i.e., non-redundant, LIO flight system de- to as the NASA/DoD/DoE/Industry Laser Ordnance
sign approach was not only considered acceptable but Team. For assuring success, it was determined at pro-
even essential to prove operational functionality. While gram onset that the conduct of critical program reviews,
the quantity of test hardware was statistically low, the namely, the Initiation, Design, Full Scale Pre-test
test program was carefully designed to minimize techni- Readiness, and Flight Readiness Reviews, were all es-
cal risks. Operational goals that closely simulate this sential.
technology's claimed functional advantages of electro-
Implementation Instrument. A new contractual instru-
magnetic insensitivity were to be incorporated into the
ment was the most important ingredient at the begin-
program's operations, within program safety and cost
ning of this program to accommodate a rapid program
constraints. It was not intended to address all LJ.O-
implementation. To that end NASA announced in the
related issues, but within programmatic limitations the
second NASA Pyrotechnics Systems Workshop, held at
results were to have widest applicability. Finally, we
the Sandia National Laboratory on February 7, 1994
desired to rapidly transfer results to industry.
that this program was being proposed for consideration.
In view of the above program philosophy, the program The NASA Cooperative Agreement for Profit Making
staff established the following flight test objectives. Organizations was the instrument planned, the first such
The primary LOSRD Program success criteria were having been implemented successfully for the initial use
based upon (1) the safe use of LIO and (2) the operation of LIO on-board the Pegasus® launch vehicle.
of laser actuated detonators from the ground and in flight
Among the criteria for a Cooperative Agreement for
for the most critical flight vehicle applications. As a
Profit Making Organizations is government participa-
American Institute of Aeronautics and Astronautics
tion. Each participating organization brings forth an
Table 1. Program Milestones.
element for the common benefit of the technology to
1994
industry. Both parties agree to share an active role in
Feb. 7 Announcement of intent: NASA Pyro-
implementing project objectives. These are non-fee
technic Systems Workshop
bearing instruments. The program content must have
general relevance and benefit to industry. No hardware April 26 Receipt of unsolicited proposal to change
is delivered to NASA. the original cooperative agreement
May 3 Evaluation team established; proposal
This is the manner by which the LOSRD Program was
review initiated
accomplished. From the government's contributions,
June 14 Change request package submitted to
NASA Headquarters provided the overall program man-
procurement
agement. NASA's Wallops Flight Facility provided the
Aug. 19 Negotiations/clarifications. Change ap-
project management for this flight mission, as well as
proved by HQS.
the pay load team and project support personnel. WFF's
responsibilities included the design, fabrication, assem- Aug. 30 WFF Project Initiation Meeting
bly, integration, and testing of the payload systems. Nov. 14 Design Review
WFF provided the instrumentation for facilitating and 1995
verifying LIO functions, such as the onboard monitors, Jan. 3 Flight hardware build-up complete
power supplies, timing devices, and transmitter as well Jan. 18 Qualification testing complete
as the data receiving station, RF tracking and photo- Feb. 7 Pre-Test Readiness Review; Full-Scale
graphic coverage, the preflight performance analysis, System Test conducted
and analysis of test and flight data. NASA provided a
March 7 Mission Readiness Review
dedicated Nike-Orion launch vehicle and facilities, a
March 15 Launch
mounting platform for installation of the laser ordnance
Apr. 25-26 Laser User-technologist Workshop at
hardware, the vehicle flight dynamics analysis, 3-axis
GSFC
accelerometer, FM-FM transmitter, flight system build,
flight system integration, and the flight performance
analysis. WFF specified standards of workmanship and
LOSRD Cost
safety throughout all phases of the project. They coor-
Demonstrations of this type can be inexpensive to im-
dinated and conducted all launch operations. Finally,
plement using the above management/team approach.
Headquarters was to assure the rapid transfer of the tech-
The cost of the agreement was S134.5K. The role of
nology.
the WFF was key to implementing a low cost, highly
The Ensign-Bickford Company (EBCo) provided the reliable program quickly.
laser ordnance system and the support necessary for in-
tegration of the ordnance onto the sounding rocket 2.0 PROGRAM DESIGN AND TEST
which was based upon an extensive base in those areas. APPROACH
This included the laser firing unit, the fiber optic cable,
connectors, detonators, initiators, and shaped charge. Requirements
The Ensign-Bickford Company developed the firing Whereas programs normally seek minimum stress lev-
system's ground support checkout equipment, provided els, the requirements for this program were the oppo-
the ordnance procedures, and installed the ordnance. The site. Subjection of the LIO hardware to the greatest
ordnance quality program was at their discretion. flight stress was the goal. Thus, the minimum payload
NASA requested that the LIO be demonstrated to meet mass, consistent with flight stability, was designed. To
vehicle operational and safety requirements and that The achieve the maximum dynamic flight environmental
Ensign-Bickford Company should establish ordnance exposure to demonstrate LIO FTS, thrust termination
test requirements. Testing was to be performed to dem- was planned at or near Max-Q. The minimum neces-
onstrate compatibility and functional performance of sary flight instrumentation to provide useful engineer-
laser initiation with the current motor ignition system. ing data was requested. Once the design was completed
and determined to meet the mission objectives, no new
Schedule Milestones.
goals or "improvements" that would cause hardware
The target time for the conduct of this demonstration changes were permitted unless safety were affected.
was to be within approximately 6 months from go-
Safety and Mission Assurance Considerations
ahead, with a report to follow. The actual time from
completion of the program initiation meeting to launch For safety, a three failure tolerant (man-safe) design of
was only 6 months, two weeks. The total time from critical safety-related events was established.
receipt of The Ensign-Bickford Company's unsolicited The lowest program cost could only be accomplished by
proposal until launch was 10 months, 19 days. Table 1 the use of off-the-shelf hardware to the fullest, even if a
presents specific program milestones. non-optimal LIO design resulted. Refer to Table 8 for
the list of actual hardware and pedigree. Simply stated,
we used what in essence amounted to a "standard con-
cept" in order to accomplish the program quickly,
American Institute of Aeronautics and Astronautics
cheaply, reliably, and safely: Previously flown hardware
Testing
was either used within prior operational constraints, or
it was retested. A typical programmatic test approach was selected; that
is, the normal acceptance and qualification testing was
It was important to keep the vehicle simple; and, conse-
performed. The greatest quantity of hardware consistent
quently, no RF command up-link was requested. An all
with program funding constraints was tested to the most
solid state firing/safing system was to be utilized, i.e.,
demanding environmental test levels necessary for meet-
no fiber optic line barriers were allowed, requiring an
ing mission success. Rigorous component level testing
electronic safe and arm system.
was first conducted on new or modified designs prior to
Hardware redundancy was excluded except for the altitude their incorporation into the payload. The complete pay-
switch, a mechanical device safety interlock, which load was then tested for the new flight environment in
safes the system from inadvertent ignition at altitudes its final configuration. Critical system functions were
below 5,000 feet. Thus, from that non-redundant hard- verified during and after environmental testing. Tests
ware approach, initiation events would unquestionably were performed on new or on modified designs and on
verify LIO operation. This single string design was prior hardware designs used in a new operational envi-
intended to be a high risk design approach, but one with ronment. Component tests, integration tests, and a
risk mitigated by a strong qualification test program, full-scale system test were all conducted. The full-scale
including an integrated ground systems test. integrated ground system test was critical to the pro-
gram, the success of which was established as a launch
Design
constraint. Testing was accomplished at various loca-
The Nike-Orion was selected as a suitable high perform-
tions: The Ensign-Bickford Company, Wallops Flight
ance launch vehicle capable of meeting the program's Facility, Johnson Space Center, and the Orbital Sci-
schedule. The redundant motor initiators normally used
ences Corporation.
were changed to a single initiator. High quality parts
The wisdom of the test approach was verified when the
were used. Key flight events, were monitored via 19
only failure which occurred was due to an inadequate
channels of flight data: 3-axis accelerometer. commands,
solder joint that failed in the laser firing unit during
arm status, battery current, and laser voltage. That in-
vibration acceptance testing. This failure was detectable
strumentation was considered sufficient for safety status
only by monitoring performance status during environ-
determinations, engineering evaluations, and mission
mental testing. Had it remained undetected, the flight
success information. Backup flight data, if necessary,
termination event would not have occurred.
were to have been provided by camera and radar cover-
age.
3.0 VEHICLE DESCRIPTION
While there were 23 critical single failure points that
could cause loss of mission, the safety reviews identi- Launch Vehicle
fied no single safety critical failure. The key safety A standard Nike-Orion launch vehicle, figure 1, was
SUP -JOINT S EPARATION
TELEMETRY/POWER SECTION
SYSTEM
/T\ L LIO SECTION
\
t \
16.50" 14.0"
DRAG RING
I PAYLOAD I
-NIKE- - ORION-
'SECTIONS '
0.0" 136.73" 143.29" 247.99" 287.99" 360.69"
Figure 1. Nike-Orion launch vehicle configuration.
features of verifiable, three-inhibit levels for FTS acti- selected to perform this mission's objectives. It offered
vation and motor ignition system were determined to the high accelerations that were required to meet mis-
have been met when reviewed by the sion objectives. Transient ignition accelerations during
NASA/DoD/DoE/Industry Laser Ordnance Team. Many lift-off can be on the order of ±50 g's, with a second
of these 23 single failure points were necessary in this peak of approximately ±30 g's occurring at second stage
ordnance system to maintain low costs that resulted ignition.
from not requalifying hardware. Hence, that number
The Nike-Orion launch is a two stage, unguided. spin
would be less in a flight system designed specifically
stabilized rocket which uses Army surplus motors.
for LIO.
This vehicle has been a mainstay of the NASA Sound-
ing Rocket Program for many years, and the LIO flight
project was its 106th sounding rocket mission. The
American Institute of Aeronautics and Astronautics
Nike-Orion can carry a 150 pound payload to 190 kilo-
The XM 22E8 Orion (Hawk) dual thrust second stage
meters, or a 450 pound payload to 90 kilometers at an
sustainer was first produced in 1960 for Army surface-
85° launch elevation (Figure 2). Care had to be exer-
to-air tactical missiles applications. The center-type
cised with the use of higher elevations to assure ocean
ignitor contains A1/KC1O and 5A black powder. The
impact of the debris after FTS activation. 4
two types of composite, solid propellant grains, cast
into a concentric-cloverleaf configuration, provide two
distinct thrusting modes. The thrust-time history of the
Launch Elevation LOSRD Mission Design motor is one that becomes highly regressive at ap-
proximately 5.5 seconds into the bum at a 60°F refer-
ence temperature. The first mode is the "boost mode"
200 Ib (90.7 kf> which operates at a chamber pressure of 1070 psig to
250lb(M3.4k|) produce an average thrust of 13,000 pounds at a specific
impulse of 235 Ibf-sec/lbm. This is followed immedi-
300 Ib (136.1 k|)
J5«IMI58.8kj> ately by the "sustain mode," which has a 26 second
40016(181 4k|)
bum duration, at a very low chamber pressure of 190
450 Ib (204.1 k») Plytad Weilta psig, producing 1740 pounds of average thrust at a spe-
cific impulse of 204 Ibf-sec/lbm. The desire to reduce
the debris pattern, by maximizing the destruction of the
payload, prompted the decision to initiate the FTS event
3.5 seconds after Orion ignition, during the higher mo-
tor chamber pressure of the boost phase. The motor
utilizes 4 fins.
1
200
Ensign-Bickford Through Bulkhead Initiators (TBI) re-
Impact. Range (km)
placed the Holex 3300 Ignition cartridges which are
Figure 2. Nike-Orion launch vehicle performance. typically used for the ignition of both motors. A slip
joint separation system provided the means for staging.
The LJO mission was launched from the Pad 2 ARC
The 22.5° fiber-glass nose-cone had a ballast weight of
(Atlantic Research Corporation) launcher on the NASA
approximately 60 Ibs. added to provide vehicle flight
Wallops Island Launch Facility at a 78° elevation and a
stability.
145° azimuth with a 185.75 pound payload. As can be
noted from figure 2, this mission was designed to ex-
Payload Design. Manufacture, and Testing
tract the highest level of performance from this launch
The payload hardware was located on 4 decks. Those
vehicle as intended by mission objectives.
were mounted in two separate payload sections called
The M88 M5-E1 Nike Hercules first stage booster has the Telemetry/power Section and the Laser Initiated
been in operational use for more than 30 years. The Ordnance, LJO Section, each 20 inches in length and 14
head end-type ignitor contains A2 black powder, and the inches in diameter. Those two sections plus a nose
double-base, solid propellant grain has a concentric ring cone with ballast comprised the payload assembly. The
configuration. Burning for a nominal 3.2 seconds, the payload was designed, fabricated, assembled, wired, and
Nike motor provides an average rated thrust of 42,782 tested at the Wallops Flight Facility. In late August
Ibf. and specific impulse of 195 Ibf-sec/lbm at 60° F. 1994, the WFF staff began to design the payload.
When used as a first stage booster, the Nike motor has
The Telemetry/power Section housed NASA's flight
three 4.8 ft2 Nike-Ajax fins. A 69 in2 concentric drag
support equipment, e.g., batteries, timer, accelerome-
plate was added for this mission to maximize separation
ters, telemetry system, etc. on decks #1 and #2, figure
from the Orion motor prior to second stage ignition.
3.
American Institute of Aeronautics and Astronautics
Figure 3. LOSRD Payload, Telemetry/power Section.
Power for the LDFU was provided from two on-board To accomplish the primary goal of exposing LIO hard-
silver-zinc batteries on deck #1. Safe and arm com- ware to the most extreme dynamic environment, a
mands for the LDFU were provided by a control panel minimum flight vehicle weight maximized the effect of
in the blockhouse. The pre-programmed Multi-function this vehicle's nominally turbulent performance charac-
Timer, located on deck #1, provided the fire commands teristics. Flight system design complexity was mini-
for the two stage events. The timer was started by a mized, and onboard instrumentation was limited to
lanyard as the rocket cleared the launcher. The timer monitoring of only those functions most critical to the
provided a fire command at T+8.0 seconds for stage 2 verification of programmatic success and as needed for
ignition and a second fire command at T+l 1.5 seconds safety. All design work was completed in early No-
for stage 2 flight termination. An FM-FM telemetry vember. The LIO Flight Vehicle Design Review was
system, deck #2, operating at 2269.5 MHz, provided held at WFF on November 14, 1994. Fabrication of
accelerometer and LIO system status data during flight. the hardware began immediately thereafter, and this
hardware development phase of the program was com-
Two decks were provided for installation of the laser
pleted by the end of January 1995.
initiated ordnance payload equipment, decks #3 and #4
which respectively contained the ordnance and the The payload's LIO Section, which housed the bulk of
LDFU laser firing system. the laser initiated ordnance system, including most of
the devices required to initiate second stage ignition and
The breakdown of the payload weight assembly is
flight termination, was located at the aft end of the pay-
shown in Table 2.
load and was mated directly to the forward end of the
Table 2. Payload weight summary.
Orion motor. A standard WFF electronic safe and arm
Assembly Weight. Ibs. (Ledex) switch was incorporated into the flight system
LIO Components/ordnance 12 as one of three safety inhibits.
Electrical, electronic, telemetry, 32 The Telemetry/power Section was located just forward
power, wiring, instrumentation of the LIO Section. The WFF wraparound S-band te-
Payload structure 55.3 lemetry antenna was mounted externally on this section,
Nose-cone 28 figure 5, as was the lanyard-type lift off switch which
Ballast 58.5 would stan the second stage ignition and FTS timer at
Total 185.8 first motion. Mounted inside the cylindrical structure
were decks #1 and #2, to which most of the WFF pay-
load components were attached. These included the bat-
tery pack power supplies and the timers, as well as the
American Institute of Aeronautics and Astronautics
redundant altitude switches which would prevent a pre-
mature second stage ordnance ignition in the event of an
accidental lanyard deploy prior to the launch. The 2
watt TM downlink RF transmitter, also mounted in this
section, provided 19 FM/FM channels of data. This
was used to transmit monitored data for flight events
which were used to verify mission success in lieu of
sufficient corroborating evidence from optical and radar
tracking. The data to be provided by this on-board sys-
,«-*••»-"*
tem included the output of a 3-axis accelerometer, laser ; ig
arming status, as well as power supply current and volt- Orion Motor Ntk« Motor
age data. Payload power was supplied by 2 silver zinc
batteries. An 11-degree, 14-inch diameter asbestos phe-
Orion Motor Head End DestmctTest
nolic nose cone was mounted at the forward end of the
telemetry section of the payload.
Figure 4. Integrated ground test set-up.
Combined System Level Testing
actuation could be verified even in the event of an un-
The laser hardware testing at the component level began
likely subsequent disruption of the explosive train dur-
in mid January and was completed by the end of the
ing the flight test. During this ground test, the onboard
month. The Ensign-Bickford and NASA/WFF payload
accelerometer data clearly indicated a ringing of the pay-
systems were then brought together in an integrated
load structure each time a LID functioned. Refer to
assembly for the Full-scale Integrated Ground System
figure 20 for a description of the dynamics (-1200 Hz
Test which was performed successfully on February 7,
ringing) of the FTS event in flight. Subsequent com-
1995. The overall test set-up is shown in figure 4.
parisons with flight test data from other programs con-
This test utilized the actual flight hardware as much as firmed that, in lieu of other corroborating evidence, this
possible. Expended Nike and Orion motor casings were method would be sufficient for verification of any suc-
substituted for the cessful LID event.
rocket motors. The
After the inte-
shaped charge and
grated ground sys-
TBI's were from
tem test, both
the same manufac-
payload system
turing lots as the
sections and the
flight items. The
complete flight
payload was sus-
assembly were
pended in order to
subjected to the
limit dampening of
comprehensive
structural harmon-
environmental
ics during firing of
testing procedures
the Laser Initiated
which have been
Detonators (LID).
developed by the
The overall ground
NASA Sounding
test configuration
Rocket Program
inclusive of the
as the standard
payload is shown ke Motor payload test series
in figure 5.
(Table 3').
ssmmta temetryfi
The program suc-
afljewgr . ir
cess criteria was
-Section • ,,
LID detonation. (•;iS •; •' «,
Hence, an impor- S-Band ,-;'
tant objective of ©term a
this test was the '
Laser £
acquisition of the
Ordinance
LID firing signa-
i Section >,' • /
ture within the
"' ' '
environment of the
payload flight con- a• •'y....-i--o .."a•nmdny nAprg-syi^s- e•""<m&-" f
figuration. This
was necessary in
Figure 5. Suspension of payload assembly for the integrated ground test.
order to ensure that
a successful LID
8
American Institute of Aeronautics and Astronautics
with the confidence of having seen this payload survive
Table 3. Pay load Test Series.
such a rigorous ordeal, we proceeded to the Pad 2 launch
Test Predicted flight Qualification test
rail for final vehicle staging with the actual flight mo-
levels levels
tors, and we prepared for the launch of the LIO demon-
Thump test Not applicable 55 Hz natural fre-
stration flight.
quency
Sine sweep Not applicable 7.3 in/s 5-89 Hz, Predicted Flight Performance
vibration 10.5g 89-800 Hz,
IS.Og 800-2000 Hz As mentioned, the vehicle mass had been minimized to
Random 8g RMS, 11.5 20g RMS, 20 expose LIO to the worst case flight dynamic environ-
vibration seconds sec/axis, 3 axes ment. The specific loads which that philosophy pro-
Bend test 58,000 in-lbs 48,000 in-lbs at duced, one may clearly observe from the predicted accel-
payload base at a 3° erations for the LIO hardware which are shown in figure
angle of attack 6.
During this payload test series, a problem was encoun-
tered with the telemetry system which resulted in the 23.0 -
change-out of the telemetry calibrator.
Those tests were conducted at the telemetry ground sta-
20.0
tion and the environmental test lab of the WFF Techni-
cal Support Branch. Before they could begin, a delay of
several days was required for resolving a noise problem
150
in the on-board telemetry system. First detected during
the full scale integrated ground test, it was determined to
be the result of faulty signal conditioning circuit. The
10.0
problem was corrected, and the payload test regime
commenced with the standard telemetry system checkout
No Flight Tenniiauoo Event
which was completed without incident.
After post-ground test refurbishment, such as ordnance
replacement, the WFF and Ensign-Bickford payload
systems were integrated as planned. Pre-vibration se-
quence testing verified that all systems were performing
to the overall payload design specifications. Systems
which were specifically verified to be fully functional
included those of the prelaunch power and arm com-
mands, the launch abort disarming sequence, the timed
flight events, as well as the power status and the current
consumption during all events. All payload hardware
demonstrated nominal performance, including the on-
board timers and relays, and the on-board monitors for -15.0 -
00 100 30.0
each component or system.
Once we had verified that all payload systems were fully Figure 6. Vehicle acceleration time history.
operational, the complete assembled payload in its
The flight termination sequence was originally planned
flight configuration was then subjected to the environ-
at max Q which occurs on this vehicle in this mission
mental testing requirements. In this sequence of tests
at 12.5 seconds. Those loads on this vehicle, when
the payload physical properties were measured and dy-
flown to this mission profile, produce the dynamic
namically spin balanced. Then it was subjected to
pressure (Q) time history as shown in Figure 7.
stresses comparable to the aerodynamic loads predicted
to occur in the flight environment. Those loads were
applied in the form of a bend test, in which the payload
was flexed to simulate the worst case aeroelastic reac-
tion of that particular payload configuration to the op-
posing forces of thrust and drag; and a vibration se-
quence was applied which simulates or exceeds the type
of random agitation that normally occurs during motor
burn. Payload functions were monitored during the
vibration tests, and experiment functions were also
monitored before and after each of them. Finally, after
all environmental testing had been completed, a repeat
of the operational sequence test verified that the payload
systems and hardware had indeed survived intact. Armed
American Institute of Aeronautics and Astronautics
Table 5. Vehicle mass summary.
Sub-assembly. Weight, Ibs.
Payload 185.8
Orion 931.7
Nike plus Interstage adapter/drag plate 1320.9
Total 2438.4
4.0 LASER INITIATED ORDNANCE AND
FIRING CIRCUIT
Design/pedigree
The design of the ordnance system was based on a laser
diode ordnance system previously designed for the Naval
Research Laboratory's Advanced Release Techniques
;3000 experiment2. The flight ordnance system for the Nike-
Orion launch vehicle comprised a two-output Laser Di-
ode Firing Unit (LDFU), a two-channel Energy Transfer
System (ETS), and Laser Initiated Detonators (LID).
This two-channel system is identical to the hardware
previously qualified for the Pegasus® laser ordnance
flight experiment3. The LDFU is an electronic safe and
arm device containing high power laser diodes controlled
by solid state transistor switches. The ETS is a two-
0.0 10.0 20.0 30.0 40.0 30.0 channel fiber optic cable assembly used to transfer the
laser light from the LDFU to the LID. The LID is an
Figure 7. Vehicle dynamic pressure.
all-secondary detonator with an output which duplicates
Key flight parameters for the flight events are provided a standard 1 grain HNS detonating cord output.
in table 4.
Ordnance system layout. The LIO ordnance system
initiated three key mission events:
Table 4. Predicted flight dynamic environment. • stage 1 ignition,
Event Time, Altitude, Range, ft Velocity, • stage 2 ignition, and
seconds ft fps • stage 2 thrust termination.
Nike 0 0 0 0 Stage 1 was ignited by the LDFU mounted on the back
Ignition
side of the launcher. (See figure 14.) The stage 2
Nike 3.54 3699 840 1964
events were initiated from the second LDFU located in
bum out
the Telemetry/power Section of stage 2. As mentioned,
Orion 8.00 11332 2725 1595
the ordnance system was specifically designed to be
Ignition
single string with no redundancy. While this provided a
FTS 11.5 19398 4800 3223
riskier mission than using redundancy, it provided an
unqualified demonstration of success.
But later, near the date of launch, there was a concern by Stage 1 ordnance system. A schematic of the Nike
the range safety office whether the vehicle might not stage 1 ordnance system is shown in figure 8. It com-
remain sufficiently intact that flight debris genera- prised an LDFU controlled from the blockhouse, ETS,
tion/dispersions would be less than desired, so the ROM LID, manifold, self-separating Flexible Confined Deto-
timing chip processor was reprogrammed only two days nating Cord Assembly (FCDCA), and Through Bulk-
before launch in order to relieve motor chamber pressure head Initiator (TBI).
at maximum chamber pressure, this motor having a
highly regressive grain design, which corresponded to
the flight time of T = 11.5 seconds. This modification
set a record for the completion of such a significant
change at a time so near launch!
The launch vehicle used standard flight control hardware
consisting of a three-fin Nike design and four-fin Orion
design (figure 1) which spin stabilized the vehicle at 7
revolutions per second. The components and their
weights used for mission performance analysis are
summarized in table 5.
10
American Institute of Aeronautics and Astronautics