CALIFORNIA STATE UNIVERSITY, NORTHRIDGE
LABORATORY SIMULATED RELIABILITY TESTING
• II·
OF A NAVY AIR LAUNCHED MISSILE
A thesis submitted in partial satisfaction of the
requirements for the degree of Master of Science of Engineering
by
Robert Donald Schilken
~·
July, 1976
The thesis of Robert Donald Schilken is approved:
California State University, Northridge
June 4, 1976
ii
TABLE OF CONTENTS
Approval
············~~~-·"·····~~~~'·"··
..
Table of Contents
List of Illustrations
...
.
'
. ...
ii
. . iii
...
iv
Definition of Terms .
vi
Abstract
. vii
CHAPTER I - Introduction to SIDEWINDER AIM 9L Test and
Evaluation . . . . . . . . . . . . . . . .
1
CHAPTER II - Historical Development of Air-To-Air Missile
Reliability Testing . . . . . . . . . . .
4
CHAPTER III -SIDEWINDER Thermoacoustic Captive Flight
Simulation . . . .
10
Varying Vibration
10
A.
Determination of Varying Vibration
B.
Implemenation of Varying Vibration
....
20
....
En vi ronmenta 1 Controls
....
Failure Determination . . . . . .
GCS Test Set Design . . .
Thermal Conditioning
.
22
23
27
27
CHAPTER IV - Results of AIM 9L Reliability Test.
.
. .......
.....
....
CHAPTER V - Conclusions .
References
Appendix A:
33
....
AIM 9L Guidance and Control Section Functional
Test Set . . . . . . . . . . . . . . . . . . .
iii
11
39
40
41
LIST OF ILLUSTRATIONS
1.
SIDEVJINDER Missile Exploded View. . . • . . . . . . . . .
2
2.
Acoustic Chamber for Simulating Aerodynamic Vibration
8
3.
AIM 9L Expected Dynamic Pressure During Missile
Lifetime . . . .
13
Table of Average Dynamic Pressures and Overall
Sound Pressure Level for Simulating SIDEWINDER
Capti~e Flight Lifetime
14
One-third Octave Band Spectrum for Acoustic
Testing of Assembled Externally Carried Aircraft Stores
16
Overall Sound Pressure Level Required to Simulate
Dynamic Pressures as Determined by an Instrumented
Captively Flown Missile and by MIL-STD-L810C
Method 515.2 . . . . . . . . . . . . . . . .
17
SIDEWINDER Continously Repeating Acoustic Cycle
for Simulating Operational Vibration . . . . . . .
19
Thermal Shroud used in SIDEWINDER Thermo-Acoustic
Simulation
24
SIDEWINDER Laboratory Reliability Testing Daily
Thermal Cycle . . . . . . . . . . . . . . . . . . .
25
AIM 9L Target Detector Functional Test Set Used
for Thermo-Acoustic Testing
30
Guidance and Control Section and Target Detector
Functional Test Sets used for Thermo-Acoutic Testing
31
SIDEWINDER Daily Environmental and Testing Cycles
for Laboratory Simulated Reliability Testing . . .
32
AIM 9L Navy Technical Evaluation Reliability Test
Failure History
35
4.
5.
6.
7.
8.
9.
10.
11.
12.
13.
iv
14.
AIM 9L Laboratory Reliability Testing Constant
Confidence Sequential Test Score Sheet . . .
36
AIM 9L Navy Technical Evaluation Laboratory
Reliability Test Failure Descriptions
38
16.
SIDEWINDER GCS Functional Testing Setup
42
17.
Simplified Block Diagram of SIDEWINDER Guidance
and Control Section
44
18.
Approaches to SIDEWINDER Simulated Target Motion .
50
19.
SIDEWINDER Simulated Target Motion used for ThermoAcoustic Reliability Testing . . .. • • • . . . . .
52
SIDEWINDER Guidance and Control Section Functional
Test Set Main Frame
54
21.
AIM 9L GCS Functional Test Log Sheet
55
22.
SIDEWINDER Guidance and Control Section Functional
Test Set Wiring and Block Diagram . . . . . . . .
58
SIDEWINDER Guidance and Control Section Functional
Test Set Main Frame Rear View . . . . . . . . . .
65
15.
20.
23.
v
DEFINITION OF TERMS
CAPTIVE CARRY - Used in the SIDEWINDER program to designate the
the condition when a missile is captively attached
to a non-flying aircraft.
CAPTIVE FLIGHT - A condition where a missile is captively attached to
a flying aircraft.
CANARD - Aerodynamic control surfaces of a missile placed forward of
the center of gravity.
CHARGE AMPLIFIER - A device to convert the output of a capacitive
accelerometer into a voltage.
FLIGHT TESTING- Testing a missile using real flight rather than
laboratory simulations
FREE FLIGHT - The condition after a missile is launched and is in
control of itself.
GCS - Guidance and Control Section.
PATE - Production Acceptance Test and Evaluation Division of PMTC •
.
PMTC - Pacific Missile Test Center, Point Mugu, California.
MTBF - Mean Time Before Failure.
NTE - Navy Technical Evaluation.
NWC- Naval Weapons Center, China Lake, California.
RATE TABLE - A mechanical device which rotates the GCS in the
horizontal plane at a controlled rate for measuring
GCS parameters.
TD - Target Detector.
vi
ABSTRACT
This report deals with the design, development, implementation
and results of a laboratory simulated reliability test of the
Navy AIM 9L (SIDEWINDER series) Air-To-Air Missile.
General facts
of the missile and of Navy Test and Evaluation are presented
followed by a brief history of laboratory environmental reliability
testing methods.
m~ntal
The termo-acoustic method of simulating environ-
conditions of missile captive flight is developed, and the
use of varying instead of fixed level vibration, unique to the
SIDEWINDER reliability test, is described.
The approach to
determine and implement the acoustically induced vibration and the
shrouded thermal conditioning is presented.
The subject of failure
determination. is discussed and the requirements for functional
testing of the missile•s operation are presented.
A detailed
description is included of the Guidance and Control Section
Functional Test Set which was designed around the unique constraints
of the thermo-acoustic reliability testing method.
The results
of the AIM 9L laboratory reliability test are given and conclusions
as to the effectiveness of the test are drawn.
The report
emphasizes the decision and inovations required to approach the
problem of providing an accurate laboratory simulation of real
j
'i
life events at minimal cost.
vii
CHAPTER I
INTRODUCTION TO SIDEWINDER AIM 9L TEST AND EVALUATION
The AIM 9L is the latest in the series of SIDEWINDER missiles
that have been developed by the Naval Weapons Center at China Lake,
California, for Naval Fleet Air-to-Air defense.
The SIDEWINDER
series of missiles, which was first conceived in the late 1940's, is
designed around a five inch diameter, twelve foot long, canard
controlled configuration as shown in Figure 1.
The four sections
of the missile consist of a solid propellent motor to provide a
burst thrust, an expanding rod warhead to disable the target (aircraft), a Target Detector (TO) to determine the time at which to
detonate the warhead, and a Guidance and Control Section (GCS) to
track the target and provide the intelligence to guide the missile
to the target.
Due to its complexity, the GCS is the most unreliable
section of the system, and is therefore the section which received
most of the attention in the AIM 9L Navy Technical Evaluation (NTE)
performed in fiscal year 1975 by the Pacific Missile Test Center
(PMTC) at Point Mugu, California.
The NTE is part of the standard testing of a weapon system
performed on newly developed weapon systems prior to being approved
for full production and introduction to the fleet.
The NTE includes
such tests as laboratory flight simulation, missile handling (by
personnel and equipment), compatibility assessments, flight tests,
countermeasure experiments, lethality analysis and reliability tests.
The GCS and TO reliability portion of the AIM 9L NTE was assigned
to the Production Acceptance Test and Evaluation (PATE) division,
1
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3
whose mission is to plan, develop and implement testing programs on
production (fully developed) missile systems.
The reliability test
was performed in the Environmental Simulation Laboratory where the
physical conditions of captive flight were simulated on four missiles
simultaneously.
This report describes the development, implementation
and results of that test with specific emphasis on the physical
methods used to generate the environment.
CHAPTER II
HISTORICAL DEVELOPMENT OF AIR-TO-AIR MISSILE RELIABILITY TEST
When Test and Evaluation (T & E) of Air-to-Air missiles was
first performed at PMTC (formerly Naval Missile Center) in the
1940's and 1950's, most testing consisted primarily of flying a
missile at a target within the constraints of the design performance
and basing the evaluation on this limited senario performance.
However, as missile systems progressed, the performance envelope
(bounds of parameters in which missiles will perform) became larger,
requiring more data to adequately assess the total performance
bounds.
Now, with the increased cost of flight testing (from $1000
to $3000 per hour) as well as the increased cost of the missiles
themselves, it has become necessary to conduct a significant part
of the T & E ih laboratory situations where the horrendous costs of
flight testing is not incurred and where the missiles are not
necessarily expended.
The use of laboratory testing is most significant today in the
area of reliability testing.
Missiles are now being produced with
such reliability that Mean Times Between Failures (MTBF) of the
systems are approaching several hundreds of hours.
In order to
evaluate a high reliability missile system, either a large sample
of the missiles must be operated to a few failures or a small
\i
I
sample must be operated to many failures.
With either approach,
the number of missile operating hours required to reach a meaningful
decision can be in the thousands.
Therefore, in the past decade
much effort has been expended towards simulating in the laboratory
4
5
at low costs the conditions of missile flight.
Air-to-Air missile system reliability, although frequently
considered as one entity, is basically a measure of the missile's
ability to (1) be captively carried on a flying aircraft (defined
as Captive Flight) for many hours, loaded and unloaded, handled
and stored; (2) be launched and fly successfully for a short
interval of time to intercept the target (defined as Free Flight).
Each part of this requirement has its own degree of environmental
severity, operational demands and failure modes.
would be desirable to evaluate each phase
Although it
individually during every
laboratory test, this is seldom done because parts of the missile
which operate during the free flight portion of the missile's
lifetime are designed to operate only for minute$, and will fail
if operated
r~peatedly
throughout a testing program.
Therefore,
most environmental simulations engage only a fixed level of
environmental severity, duplicating only the free flight, or only
the captive flight portion of the mission.
When environmental simulation was first being developed,
attempts were made to duplicate in the laboratory all possible
physical environmental elements, such as temperature, humidity,
altitude, vibration, acceleration, shock, etc.
It was, however,
recognized very early that the two most detrimental environmental
)
!
conditions to a weapon's performance and reliability are temperature and vibration.
The combined effect of these two elements
produce more than ninety five percent of the environmentally
induced missile failures in a modern weapon system.
6
One of the first facilities developed to simulate temperature
and vibration in the laboratory was an Agree Chamber.
This device
consisted of an electromechanical shaker of some type to induce
vibration into the missile, with a thermal insulating shell surrounding the missile into which hot or cold air was ducted to temperature
condition the item.
The vibration, usually a
si~ewave
single
frequency single axis, was then cycled on and off and the temperature cycled from extreme hot to extreme cold in a prescribed
manner which was eventually formalized into the MIL-STD-781
Environmental Testing Method.
This method of testing ltJas satisfactory to assess the ability
of the frame or mechanical structure to withstand extreme environments, but did not test the ability of the missile performance as
a whole to tolerate realistic environments.
In an effort to
produce a more realistic simulation, random frequency (noise)
vibration was used instead of single frequency vibration.
This
approach produced an environment that generated more realistic 1
correlatible with fleet use, typical failures.
However, using one
electromechanical shaker produces vibration in only a single axis.
The procedure then was to mount the test item in three mutually
perpendicular orientations in three separate time testing intervals.
This testing method although still used today in a large
portion of the environmental testing field is obviously not ideal.
A later advancement in vibration simulation incorporated the
use of three separate mechanical shakers, one in each of the three
orthagonal axis, loosely coupled (in the non-normal axis) to the
7
test specimen.
Using three separate sources of random vibration,
one for each axis produced a good simulation of missile flight
vibration at the point where the three vibration inputs were
coupled to the missile.
However, in real life the majority of the
vibration is not induced into the missile at only one point, but
rather is distributed along the entire surface of the missile in
the form of random pressure variations generated by the turbulance
in the aerodynamic boundry layer during flight.
Awareness of this
fact has directed recent thinking to the approach of using acoustic
energy to generate random pressure variations at the skin of the
missile thereby producing the desired vibration in the missile.
The now common acoustic method of inducing vibration for
simulation of aerodynamic situations utilizes a rigid reflective
walled chamber into which high energy acoustic noise is coupled
as shown in Figure 2.
The extreme noise energy required in this
simulation is usually generated by modulating a flow of low
pressure air at audio frequencies with an electrically driven
valve (air modulator).
The valve, which in its static condition
restricts approximately half of the air flow, is designed to respond
extremely quickly to an electrical signal to restrict more or less
of the air flow depending on the polarity of the driving current.
The modulator is then driven with a low frequency electric signal
\
in the range of 50 Hz to 2000 Hz creating variations in pressure
I
corresponding to the variations in the driving signal.
These
variations in pressure are then coupled to the rigid walled
acoustic chamber with a common exponential horn.
The result is
8
L0\4 PRESSURE
AIR SUPPLY
NOISE SIGNAL
~--~
DR IV ERt-----l
EXPONENTIAL HORN
~-
MODULATOR
·-·--~
~-----M~I~S~SI~L~E--~~~--~
RIGID WALLED ACOUSTIC CHAMBER
FIGURE 2 - ACOUSTIC CHAMBER FOR SIMULATING
AERODYNA~1IC
VIBRATIONS
9
a reverberant but nonresonant
~hamber
in which a portion of the
space has a spacially randomly distributed high energy pressure
vector.
The method of using acoustically induced vibration augmented
with the common thermal conditioning approach, called a ThermoAcoustic Simulation, is now being accepted and implemented as one
of the most accurate methods of environmental flight simulation,
and is the method chosen for use in the SIDEWINDER captive flight
reliability simulation.
CHAPTER III
SIDEWINDER THERMO-ACOUSTIC CAPTIVE FLIGHT SIMULATION
Varying Vibration
One of the more significant contributions of the SIDEWINDER
program to the area of reliability testing is inthe incorporation
of varying vibration.
Previous simulations of flight vibration,
whether using mechanical shakers or acoustic excitation, utilized
only one level of overall (average) vibration.
It was felt that
the rate of degradation of a missile was related to the amplitude
of the vibration (A) by the expression:
MTBF ::::::5 (A) 1/k
where k is a constant with a value of two to ten, depending on the
material and environmental conditions.
Based on this premise, reliability testing then used a fixed
level of vibration which simulated either a real time rate of
failures (normal vibration level) or a fixed accelerated rate of
failures (increased vibration level).
However, data gathered from
accelerated reliability tests showed that the types of failures
occuri ng in testing \</ere not the same types as were occurring in
real life or in nonaccelerated tests.
A report in 19741 indicated
that the types of electronic components failing due to vibration was
dependent on the amplitude of vibration as well as the duration.
1
J. C. Calkins, SPARROVJ III ~·lissile Accelerated Reliability
Testing, (Naval Missile Center, Point Mugu, California TP-74-12,
May 10, 197 4. )
10
11
Low amplitude vibration for long periods of time caused electromechanical devices to fail.
Moderate vibration (representative of
average flight conditons) caused passive electronic components to
fail, and high amplitude vibration for short periods of time caused
active electronic components (tubes) to fail.
During its captive flight lifetime, a missile will experience
a large range of vibration due to the various flight conditions
encountered.
The SIDEWINDER missile, in particular, experiences a
greatly diverse range of vibration due to the fact that part of its
operational life is accumulated with essentially no vibration,
during the time it is
11
Captively Carried
maneuvering on the ground.
11
on an aircraft that is
Therefore, in designing the SIDEWINDER
simulation an attempt was made to derive the percentages of time
the missile wo·uld experience various levels of vibration during
its lifetime.
A.
First the captive flight environment was defined.
Determination of Varying Vibration
A Hughes Aircraft Company Document 2 , specifying the expected
flight seneario earameters for the AWG-9 Weapon Control System in
the F-14 Aircraft was selected as being representative of the
flight conditons that SIDEWINDER would experience on a Navy fleet
defense aircraft.
The flight parameters given in the report,
which included missions from high altitude slow speed to low
\
2
Hughes Aircraft Company, Aerospace Gro~p, Environmental
Criteria Re ort for AN/AWG-9
on Control S stem,
SD-15304-R, April 1971 .
12
altitude high speed chase, were presented as a summary of the expected percentage of lifetime of the system at a given altitude band
and mach number (data Classified).
From this information air
density (p) and velocity (v) for each flight condition were derived
from standard tables, and flight Dydnmic Pressures (q) were determined
using the relation:
The range of q conditions determined versus the percentage of
missile lifetime is plotted in Figure 3.
For the purpose of
simplifying the simulation an average of each of these bands was
used and the lowest and highest q extremes, representing approximately four percent of the expected missile lifetime, were deleted.
The upper extreme was difficult to generate with existing facilities
and the lower extreme would actually be generated for the shore
period required in the randomness of the vibration applied during
the other levels.
Figure 4 summarizes the q conditions used to
represent SIDEWINDER captive flight as well as the overall sound
pressure level (referenced to 2 x 10- 5 newtons per square meter) of
the diffuse field within the acoustic chamber necessary to simulate
these q's.
The overall (average) sound pressure level (SPL) was
measured with a l/4 inch diameter
Bruel and Kjaer model 4310
capacitance microphone placed eight inches from the body of the
missile and a Bruel and Kjaer model 2113 One-third Octave spectrum
analyzer.
The amplitude of the SPL was determined using a vibration
instrumented captively flown SIDEWINDER (Aim AIM-9G serial number
13
DYNAMIC PRESSURE- q (lbs/ft2)
0
0
0
0
0
C>
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('I')
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1.0
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w
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('I')
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FIGURE 3 - AIM 9L EXPECTED DYNAMIC PRESSURE DURING
MISSILE LIFETIME
0
1--1
(/)
(/)
1--1
:E.:
% OF
CAPTIVE FLIGHT
LIFETIME
OVERALL ACOUSTIC
SOUND PRESSURE
LEVEL {2Xl0-4Njm2)
126
8%
139 db
141
17%
140 db
223
32%
144 db
302
10%
147 db
2%
148 db
439
5%
150 db
527
26%
151 db
DYNAMIC
PRESSURE
q (PSF)
355
.
FIGURE 4 - TABLE OF AVERAGE DYNAMIC PRESSURES AND OVERALL
SOUND PRESSURE LEVEL FOR SIMULATING SIDEWINDER
CAPTIVE FLIGHT LIFETIME.
15
1862) missile as a reference.
This missile had previously been
carried on an F-4 aircraft with controlled flight conditions, and
the vibration occurring in various parts of the missile recorded
on magnetic tape.
The missile was then placed in the acoustic
chamber and the amplitude and spectrum_of the chamber noise adjusted
to yield missile vibration, measured with the same instrumentation,
similar to that experienced in captive flight.
An Indevco model 2730
Charge Amplifier, a Spectra Scope model SD-330 real time
constant bandwidth spectrum analyzer, and a Mosely Autograph model
2DR-2A X-V Plotter were used to analyze the missile vibration from
each accelerometer.
For the purpose of being acceptable with the government
standard for acoustic testing, MIL-STD-810C method 515.2 was used
to verify the SPL's.
This method, which was emperically derived,
approximates the spectral response of the missile in flight with
a curve of the form shown in Figure 5 where:
f = 600 log(x/R + 400)
L = .2~ log(q} + 11 log(x) + 7 log(l - cos B)
and where:
x
= distance
along missile axis from nose to
point of vibration monitored
R = radius of missile
q
= dynamic
pressure being simulated
B - local nose cone angle
The results are tabulated in Figure 6, and indicated an agreement with the captive flight reference within 1/2 db.
10
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STORES
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17
DYNAMIC PRESSURE
INSTRUMENTED MISSILE
dB
MIL-STD-810C
dB
126
138.8
139
141
139.7
140
223
' 143.7
144
302
146.4
147
355
147.7
148
439
149.7
150
527
151.2
151
FIGURE 6 - OVERALL SOUND PRESSURE LEVEL REQUIRED TO
SH1ULATE
DYNA~1IC
PRESSURES AS DETERMINED
BY AN_INSTRUMENTED CAPTIVELY FLOWN MISSILE
AND BY MIL-STD-810C METHOD 515.2
18
The 11 Captive Carry 11 condition, where the fully operational
missile is captively carried on a non-flying aircraft in preparation
for or subsequent to flight was also simulated.
This condition
exists because the missile•s Infra-Red (IR) seeker is a mechanical
gyroscope gimballed in two axes which must be constrained (caged)
whenever the missile is moved to prevent the gyrp from damaging
itself by mechanically impacting the gimbal limits.
During missile
handling the gyro is constrained with a magnetic ring imbedded in
a rubber cap which is placed over the IR transparent protective
dome. ·However, once the missile is loaded on an iarcraft the
protective cap is removed and the gyro must be constrained electromechanically, requiring that the entire seeker portion of the
missile electronic guidance be energized.
average
fligh~
The portion of an
mission when this condition exists is estimated to
2/9 of the missile operational time and is simulated with no
vibration as shown in the overall SPL versus time of the repeated
acoustic cycle shown in Figure 7.
The h-Jenty-t\'Jo minute cycle of the vibration was arrived at by
considering the range of acceptable cycle times as follows.
In
order to insure that a missile under test will experience the
desired proportions of amplitude levels during the time it is
tested (time before a failure), many cycles (ten or more) should
occur in the testing time before a missile failure.
Assuming each
missile is a sample size of one from a large population with known
~1TBF,
a chi-square table, with tv1o degrees of freedom, was
consulted to give a minimum MTBF, with ninety percent confidence,
19
OVERALL SOUND PRESSURE LEVEL (db)
co
0
<:::!'"
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0
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N
0
0
,....
0
DYNAMIC PRESSURE (PSF)
FIGURE 7 - SIDEVJINDE_R CONTINUOUSLY REPEATING ACOUSTIC. CY~L~
FOR SIMULATING OPERATION VIBRATION
I
I.
I
that would occur in the sample.
This specifies the maximum acoustic
cycle time to be:
Cycle time ( 1/lO(MTBF population)('Xt 10 )/2
Cycle time ( 1/10(250)(0.105) hours
Cycle time ( 2.6 hours
The minimum cycle time should be sufficiently long to allow
the randomness of each level of vibration to average the desired
amplitude.
The lowest frequency observed in the random noise
generator used for acoustic simulation is 1 Hz, dictating that the
minimum time at each level should be:
Single Level Time ) 10(1/1 Hz)
= 10
seconds
A cycle time of twenty-two minutes was chosen out of convenience to
be the length of play of a standard tape cartridge.
It should· be noted that the concept was considered of making
the time at each level of vibration representative of the time
a missile would experience this level in a single mission (flight).
However, each mission is different and it would take a fairly
extensive study to determine a minimum number of different cycles
necessary to represent all missions.
Not only did the SIDEWINDER
program not have the resources to initiate such a study, but at
the present time there is no indication that this concept, if
implemented, would produce more valid failures or failure rates than
a single repeated acoustic cycle.
B.
Implementation of Varying Vibration
The normal laboratory method in the past of generating a signal
for producing the acoustics of a simulation has been to use an
21
expensive instrumentation quality reel-to-reel tape recorder.
However, the continuously repeated cycle necessary for the SIDEWINDER simulation suggested using some kind of a continuous loop of
tape, and the most popularly known continuous loop tape system is
the common entertainment type eight-track cartridge system.
Analysis of the requirements for the recording system used to
produce the SIDEWINDER acoustics indicated that:
(1) A signal to noise ratio of 15 db at the lowest level
recorded was required to insure nondiscoloration of the desired
frequency spectrum.
For an overall SPL range of 12 db (139 db to
15 db) the recording system must therefore have a signal to noise
ratio of 28 db or greater, referenced to 0 VU.
A standard home
entertainment cartridge recorder has a signal to noise ratio
greater than 40 db with new recording heads (and greater than 30 db
with worn heads and worn tape) and was therefore satisfactory.
(2) A record-playback system frequency response of 50 Hz
to 2000 Hz
! 3 db with distortion (total harmonic and intermodulation)
less than eleven. P.ercent required to insure a signal to noise ratio
of 15 db.
A standard home entertainment cartridge recorder was well
within this requirement.
(3) The playback reliability should have a time to
failure of at least 480 hours for the logistics of running a sixteen
hours per day test five days per week for three weeks.
Although
reliability data of entertainment recorders was not published, a
'survey of cartridge system owners indicated that these recording
systems in normal use exhibit an MTBF in excess of 500 hours.
22
• i;at a home entertainment recorder used for acoustic
···
- . :.
~
dtion should exhibit better reliability than in normal
,.,,
•·.he fact'that the unit will onlybe used in the playback
. nm at ~·fixed reliability control settings and will
four available tracks.
This excludes the
,., . :·:"Pccted by failures in the retarding amplifier,
the VU circuitry and the mechanical track
'!(• · ·
ud controls.
Also, because the home
·:,:-rs are inexpensive, one or more back up units
L:
0~sentially
eliminate down time if a failure
0corder se1~tted for the SIDEWINDER acoustic cycle signal
· :•.:·
'·.·
· ·,,
;~h>>.
,;h
1
readily available Audiovox model C-815R,
:~ed
Two
for one hundred dollars each and were modified
.' •:r:•> by removing the automatic channel advance
•·:mi sms.
~ ·c~
No recorder failures occurred in the program.
taken in the SIDEWINDER simulation to thermally
~lectronic
portions of the missile deviated from the
.,:.:.·· 11•\.:Lhod of heating and cooling the entire chamber.
. ,. . ...
A supple
'_;;:''DUd which is transparent (approximately) to acoustic
was used to duct the hot: or
.:
col~
air around only the
Control Section (GCS) and Target Detector (TO),
t
, (:-~ntr·ating the available thermal energy on the critical
~···.;:·~ •··
.nE: missile.
This allowedmore··ra·p"id:changes of missile
23
are experienced in real flight.
Thermal shocks induce a large temperature gradient in the
missile which produces uneven expansion of mechanical assemblies
causing stress that when combined with vibration produces unique
failures.
The physical arrangement utilized in the test is shown
in Figure 8.
The thermal cycle used in the simulation (during a sixteen
hour testing day) was derived from the AIM-9L Development Specification AS-3216.
Because the test was designed to simulate the
range of conditions most representative of actual flight conditions,
the extreme missile skin temperatures of
-6~F
and +l65°F, which
seldom occurs in real life (one percent of the lifetime) were not
reproduced.
Instead, the severe range of -40°F to +1400F was used,
as shown in Figure 9, maintaining each of these extremes for
approximately four hours to allow enough time for internal parts
of the missile to thermally respond.
The transition from one
extreme to the other was accomplished in twenty to forty minutes,
providing a 4.5 to 9 degree per second thermal shock.
Environmental Controls
Once the acoustic noise for simulating aerodynamic vibration
had been determined and documented with a Power Spectral Density
{PSD) plot from a spectrum analyzer, and the driving signal to
.
the chamber recorded, the chamber was recalibrated during the
test by bio lllethods.
Based on the assumption that the spectral
(frequency) distribution of the driving signal from the recording
system was correct, (1) the overall SPL of the chamber was monitored
24
FIGURE 8 - THERMAL SHROUD USED IN SIDEWINDER THERMOACOUSTIC
SIMULATION.
25
26
and (2) the vibration of the missile at any single point was
monitored with an average root-mean-square (RMS) g-reading from
a charge amplifier.
A charge amplifier is a device which converts
the output of a capacitive accelerometer to a voltage which is
proportional to the vecotr acceleration along the axis of the
measuring accelerometer.
A single point of acceleration monitoring
was adequate because the missile will always respond in the same
manner to the random pressure variations in the chamber.
The
vibration, therefore, did not require a closed loop continously
monitoring servo system, thereby simplifying operation of the
system and reducing the danger of a feedback failure allowing the
system to 11 run away 11 as is the case with mechanical shaker systems.
The thermal control system did require a closed loop servo
system that sensed when to apply hot or cold air to the missiles
by monitoring missile skin temperature with a copper-constanan
thermocouple mounted on one of the four missiles.
The flow of
thermally conditioned air to each missile was initially adjusted
to produce similar temperature responses in all of the missiles
thereby allowing any one of the missile's temperature response to
be representative of all of the missiles.
As a precaution against
the temperature controlling thermocouple failing and allowing the
system to run away, all four missile skin temperatures measured
with thermocouples were monitored periodically by the operating
personnel to detect such a failure and the thermal conditioning
unit was programed to limit the hot conditioning air temperature
1
to a maximum of +200•F.
27
Failure Determination
The purpose of the SIDEWINDER reliability test was to determine
the length of time each missile would operate properly under
simulated captive flight enviromental conditions.
Then using the
failure data from this sample the captive flight rel i abi 1ity of the
population of missiles from which this sample was chosen was
predicted using common statistical methods.
It was therefore
necessary to determine fairly accurately when the parts of the
missile failed to operate properly.
Because the commonly used test
sets which measure for absolute missile specificatiors were not
compatible with the acoustic testing method, special electronic
functional test sets were designed to determine periodically the
operational status of the GCS and TO's.
GCS Test Set
~esign
When the laboratory reliability test was originally proposed
as a part of the NTE, a very limited budget was submitted assuming
that a very simple method of functionally testing the missiles would
be used.
However, when the SIDEWINDER simulation was being designed,
it was determined that four missiles were to be tested simultaneously
in order to acquire enough data in the time alloted for testing.
In
order for four missiles to be functionally tested an elaborate
missile-to-test set interface (switching) would be required.
Because
of the effort that would have to be expended on this part of the
system it was decided that it would be worthwhile for future SIDEWINDER tests to invest at this time more effort than originally
planned into the test sets.
28
The first constraint then imposed on the test sets design
~as
the limited budget of $15,000 to define the parameters to be tested
and to design, develop and install the sets.
Constraints imposed
by the nature of the thermo-acoustic testing method and the acoustic
chamber were:
1.
The missile could have no bulky mechanical devices attached
which would alter the distribution of the vibration response of
the missile to the acoustic noise.
This required that a moving
target and stationary missile configuration be used instead of the
conventional angularly rotating missile (utilizing a large mechanical
rate table) with stationary target.
2.
The functional test sets must be located approximately one
hundred feet from the missile to isolate operating personnel from the
severe
environ~ent
of the acoustic chamber.
This requires special
buffering of the electronic signals to insure against capacitive
loading of the cable altering the missile•s performance.
Additional constraints were placed on the designs due to the possibility of nontechnical personnel being involved in the operation and
maintenance of the equipment.
This required the test sets to be
as simple as possible to operate, calibrate, and maintain.
The GCS test set operation was designed to be semiautomatic
in that all that was required was to advance a large rotary switch
for each test performed and the test set automatically controlled
the missile, the simulated target, and peripheral instruments used
to measure the parameters under test.
The GCS test set also indi-
cated with labeled lights which peripheral device to read and the
.
29
units of measurement (Hz, seconds, volts, etc.).
The test set
essentially required no calibration except that of the peripheral
measuring devices and had most of the electronics mounted on four
accessible edge-plugged printed circuit cards for easy removal and
repair.
A more detailed description of the GCS test set design is
included as Appendix A.
The TO functional test set operation was also semiautomatic in
that all that was required was for the operator to depress the start
button and wait until the test set indicated the testing was completed.
Three different data, displayed in three separate panel devices,
indicated values of critical TO parameters.
The circuit design, which
is Classified, is not included in this report.
A photograph of the
nineteen inch rack mounted TO functional test set is included in
Figure 10.
A photograph of the GCS and TO functional test sets
installed in a standard nineteen inch relay rack is shown in Figure
11.
Functional testing of the missiles was performed three times in
a sixteen hour testing day (bm consecutive eight hour shifts) as
shown in Figure 12.
The tests were synchronized with the thermal
cycle to provide operational verification after a lengthy cold soak,
after a lengthy hot soak, and during the transition from cold to hot.
This testing sequence was requested by the developer (Naval Weapons
Center) for certain requirements of their own data files.
In the
future these functional tests will be distributed more evenly
throughout the sixteen hour testing day.
30
FIGURE 10 - AIM 9L TARGET DETECTOR FUNCTIONAL TEST SET
USED FOR THERMO-ACOUSTIC TESTING
31
• 0
Oscilloscope
.~;;::::~'<''T"'~>. :,.
<>-•-tt
,..,
<"'""
""""
'-'""-"vu.•.,,_,, '''"•c•••
0
Target Detector
Test Set
~issile
FIGURE 11 - GUIDANCE AND CONTROL SECTION AND TARGET
DETECTOR FUNCTIONAL TEST SETS USED FOR
THERMO-ACOUSTIC TESTING
Selector
32
I
l
- -11\ ·- -
~
-·
1.0
...-
----
c-=--_
-]
l---~---.
>c::x::
0
--------··--I
(.!)
z
......
,-----
1-
-,
c·-"-
(/)
w
'----
1-
----
c?
---~,
~---------------,
..
~
c1
c::---,
~
"------,
_J
I
I
l
N
Il
I
c::x::
1-
z
I
w
::;;::
z
0
0:::
1-1
I
~I
>
z
w
I
!
~
co
(/)
0:::
~
0
::c
..__..
(/)
w
1-
::;;::
(/)
w
1-1
1-
1-
_J
c::x::
z
0
1-1
1(._)
~
z
~
u._
J_
~
~
'-----------
1
~---
I
+140
+70
TEMPERATURE
-40
c
0
OFF
ACOUSTICS
l5ldb
FIGURE 12 - SIDEWINDER DAILY ENVIRONMENTAL AND TESTING CYCLES
FOR LABORATORY Sn1ULATED RELIABILITY TESTING
CHAPTER IV
RESULTS OF AIM 9L RELIABILITY tEST
Four AIM 9L GCS's (serial numbers EXT-0054, EXT-0055, EXT-0057,
and EXT-0058) and TO's (serial numbers 388, 389, 402, and 404) were
subjected to the SIDEWINDER Thermo-Acoustic Captive Flight Simulation
during December of 1974 and January of 1975.
Prior to enviromental
testing each GCS and TC passed a full "rate table specification test
11
using the standard Deport level Navy test set located at the Missile
Preparation Branch (MPB), PMTC, which utilizes specialized optical
and rate table equipment.
The Gcs•s and TO's were also tested
prior to environmental testing by the special functional test sets
in the termo-acoustic chamber.
The first time the four missiles were functionally tested during
the simulation (four hours into the simulation) one of the GCS's
(serial number EXT-0057) failed to produce the Pilot Intercom
signal (Chirp).
This unexpected early failure was verified to be
a missile failure and not a test setup failure by switching missile
to test set interface cables.
The GCS was removed from the chamber
and subjected to a full rate table specification test which confirmed
the failure.
After investigation of the GCS records the failure
was considered to be a NO TEST" due to the fact that the failure
11
had been noted in the original full rate table specification test
performed by MPB.
The GCS had been approved for environmental test-
ing by MPB under the assumption that the Pilot Intercom problem
was caused by an intermittant test set malfunction.
33
The GCS should
34
have been rejected at the first detection of the intermittant, and
therefore was not considered part of the environmental test.
The GCS was returned to the Naval Weapons Center for failure
analysis and repair.
an
int~rmittant
Analysis determined the problem to be due to
electrical connection attributed to faulty soldering.
After being repaired by Naval Weapons Center, EXT-0057 again failed
during the incoming full rate table specification test due to its
inability to cage correctly.
This failure was detected before it
reached the environmental test and therefore was also not scored
in the NTE results.
The program was continued with the three remaining GCS's and
four TO's.
The first scored GSC failure occurred twenty hours into
the simulation with sixty total missile test hours accumulated as
shown in
Figu~e
13.
The second scored GSC failure occurred with
a total of one hundred and thirty four missile test hours accumulated.
The data point of two failures with a total of one hundred and
thirty four missile hours plotted on the Captive Flight Reliability
(CFR) score sheet, Figure 14, definitely lies in the reject region,
indicating that the true MTBF of the GCS is less than six hundred
and fifty hours with ninety percent confidence.
Scoring was based
on a 11 Constant Confidence Sequential Testing .. method 3 , which determines accept and reject decisions using a Chi-square distribution
with 2F or and 2F + 2 degrees of freedom respectively.
The AIM 9L
CFR score sheet was designed using a decision criteria of:
3
D. B. Meeker, Design of Constant·confidence Sequential Tests,
(Naval Missile Center, Point Mugu, California, TP-73-44, August 10,
1973).
35
i
MISSILE HOURS ACCUMULATED
FAILURE EXT-0054 EXT-0055 EXT-0057 EXT-0058
0
0
0
No Test
4
4
l
20
Failed
20
2
57
20
0
TOTAL
MISSILE
HOURS
0
0
4
No Test
0
20
60
0
Failed
57
134
Failed
4
FIGURE 13 - AIM 9L NAVY TECHNICAL EVALUATION LABORATORY
RELIABILITY TEST FAILURE HISTORY
36
6
2048
26387_
0
&
----ii------l-f:(.~::<---~--
5
G&
1593
.~\·
4
u..
0
n:
w
co
5z
1
I
I
·1------·----1-1
--+---·
0
500
1000
1500
2000
2500
TEST TINE IN TOTAL NISSILE HOURS
FIGURE 14 - Am 9L LABORATORY REL.IABILITY. TESTING CONSTANT
CONFIDENCE SEQUENTIAL TEST SCORE SHEET
37
(1} accept the missile reliability when there is ninety
percent confidence that the MTBF of the true population of missiles
(of which the tested missiles are a sample) is greater than two
hundred and fifty hours.
(2) reject the missile reliability when there is ninety
percent confidence that the MTBF of the true population of missiles
is less than six hundred and fifty hours.
The four TO's tested during the AIM 9L Thermo-Acoustic Captive
Flight Reliability simulation experienced no failures with a total
accumulated time of two hundred and twenty five hours.
The majority of failures that occurred during the testing
(shown in Figure 15) were attributed to Quality Control (QC)
problems of the manufacturing process and were not due to design
deficiencies.
Although the testing was terminated earlier than
anticipated, the results of the testing were meaningful and
conclusive.
38
MISSILE
HOURS TO
FAILURE
EXT-0057
6
FAILURE
Loss of Pilot Intercom signal due to
faulty soldering of the flex harness
to Module L-3.
Inability to cage due to faulty
soldering connection at umbilical
base block.
EXT-0055
6
EXT-0058
20
Up-down servo section failure:
Q-408 and Q-412.
Intermittent Caging due to faulty
soldering connection at umbilical
base block.
FIGURE 15 -AIM 9L NAVY TECHNICAL EVALUATION LABORATORY
RELIABILITY TEST FAILURE DESCRIPTIONS
CONCLUSIONS
The AIM 9L Laboratory Simulated Captive Flight Reliability
Test as performed during the Navy Technical Evaluation at the
Pacific Missile Test Center, Point Mugu, California was successful
in detecting and identifying deficiencies in manufacturing
process that caused the Guidance and Control Section to demonstrate
an unacceptable reliability.
The thermo-acoustic environmental
simulation method, using varying amplitude vibration and thermal
shrouding, proved to be an effective technique for simulating in
the laboratory the environmental conditions of captive flight.
The relatively inexpensive equipment utilized in the testing
provided a
~ost
effective substitute for actual flight testing.
The operation and control of the environment was simple and the
special designed functional test sets detected all missile
fa i1 ures.
Due to the successful performance of the laboratory tests
and the apparent production quality control problem, the AIM 9L
Guidance and Control Section will undergo a continuation of the
Laboratory Simulated Captive Flight Reliability Test at PMTC
before the missiles are released into the fleet.
39
REFERENCES
Reliability Tests:· Exponential
November 15, 1967.
l.
MIL-STD-781B~
2.
Calkins, J.C. SPARROW III Missile Accelerated Reliability
Testing. Naval Missile Center, Point Mugu, California
TP-74-12, May 10, 1974.
3.
Hughes Aircraft Company, Aerospace Group. Environmental
Criteria Report for AN/AWG-9(XN-3) Weapon Control System,
SD-15304-R. April 1971.
4.
MIL-STD-810C, Environmental Test Methods.
5.
AS-3216, Prime Item Prodtict Fabrication Specification for
Guidance-Control Section, AN/DSQ-29. Naval Air Systems
Command, Department of the Navy, June 1975.
6.
Meeker, D.B. Design of Constant Confidence Sequential Tests.
Naval Missile Center, TP-73-44, August 10, 1973.
40
Distribution~
March 10, 1975.
41
APPENDIX A
AIM 9L GCS FUNCTIONAL TEST SET DESIGN
The overall approach to functional testing of the AIM 9L
GCS is shown in Figure 16 .. Four GCS's are held stationary in
the environmental acoustic testing chamber while infrared targets ·
for each missile are moved simultaneously by a Target Motion
Drive.
The missile to be tested at any one time is electrically
interfaced to the GCS test set by the Missile Selector which
automatically upon selection of one missile places the other
missiles in a standby mode.
The standby mode is a fully powered
but non-tracking mode and is the operating mode the missile is
normally in when being captively carried or captively flown.
In
this mode, the Missile Selector powers the other missiles with
separate standby power so as not to effect the power of the one
missile under test.
The GCS test set(with its own power supply) controls the
one missile under test, the targets and target motion system
and four peripheral readout devices.
The GCS test set, hereafter
referred to as the test set, is a relatively inexpensive unit
designed around the constaints of a thermo-acoustic simulation
and a limited budget.
Th~
test set was not designed to be a
tight tolerance specification measuring device but rather a
simple missile operational status indicator that would detect
catastrophic failures and relative parameter changes with time.
The test set has two modes of operation.
A semiautomatic
mode is normally used and requires only that a rotary switch be
..
...,
......
CJ)
/'
c:
;:o
lTI
_..
0)
(./)
......
0
fTl
::E:
......
z
0
fTl
;:o
CJ)
n
'I
TJl.RGETS -.:.
. -.....~ ,~
. -''r:Z
..{.:.
~ ,il
~~
~
<
<._
MISSILE l
~1ISSILE
1
---,·····--·-············
OSC!LLOSCOPC
~
'
2
ljDr
VOLTMETER
\
·--·
I
TRUE RMS
(./)
VOLH~ETER
-n
c:
z
n
_,____
-I
......
EUNTER
I
I
!_
I
0
z)::>
r
GCS
TEST SET
-I
fTl
(./)
-I
......
zCJ)
(./)
fTl
-I
c:
-o
~1ISSILE
SELECTOR
_j
t
TEST SET
POWER
.l
+>o
N
I
43
advanced for each parameter to be measured.
The test set then
automatically controls the missile, targets, target motion drive
and peripheral devices, and indicates with marked lights which
peripheral device to be read and the units (HZ, seconds, volts,
etc.) of the measurement.
The manual mode of test set operation allows the semiautomatic control at any test (rotary switch position) to be
overidden with a series of toggle switches which manually select
certain function and modes of both the GCS and the test set.
This
manual mode is used for more thorough analysis of missile malfunctions and anomalies.
Determination Tests
The series of fourteen semiautomatic tests performed by
the test set·was arrived at by first considering the operation
of the missile and then determining the simplest means of checking
as much of the GCS systems as possible.
Figure 17 is a simplified
block diagram model of the basic SIDEWINDER Guidance system.
During target tracking the seeker look angle referenced to the
missile longitudinal axis is compared with the target angle to
procude an error which is amplified by the seeker loop gains
G and G52 , to produce a torque on the seeker gyro which alters
51
the seeker angle and reduces the error (ideally to zero). Any
external inputs to change the angle of the missile relative to
a fixed (free space) coordinate system will change the seeker
angle causing an error which will be processed by the same seeker
loop to change the seeker angle and reduce the error.
When the
r
'
AIRCRAFT
-n
,_.
G)
c
I TARGET!
;a
fT1
--'
""-1
I
G)
c
,_.
Vl
,_.
3:
):>
:z
-a
r
,_.
n
"'Tl
fT1
.......
0
):>
z
0
n
0
fT1
0
co
r
0
0
,_.
Vl
fT1
n
-1
.......
0
z
CANARDS
I
~
Error
):>
G)
;a
):>
3:
~B
i
-....5]
>'l
H
I
G53
/
Error
Sl
Gyro Angle ~
Gyro Reference
0
A
r
I
Seeker Angle
I
I GS2
~~;
~·
Correction Torque
0
""Tl
Vl
l
.......
0
fT1
.
-,
·- -
G
_A._.
t----·--l
I
- -Mi DIRECTION
n
z
-1
;a
IANGLE J
1·-
AERODYNAMICS
I
GG --
--AZ
!I\
t,J
_)[
'C;·::..·~A-z_-n
rjj__l
r
iTORQUEj
!TORQUE!
~I
M.OTOR
~~10TOR l .
!
GM
I
?r-·;~_j
I
1
I
GM
i -~\
!
I
__ j______ _j_
'·, ~- -- -·- - !
. \'Precess ""'~
iSignal 1 ."
-- -- ... -
i
r
I
I
i
1000 PSI
AIR SUPPLY
···----~----·----'
!
I
::E:
.......
z
CJ
fT1
;a
.
[
Tr~ck~ng-· ~
1. _: _
j
~~
Caged
j_ ---
LAUNCHER
-
.,J::.
:.j::o
45
missile is launched and is guiding to a target the error is also
processed by the guidance loop (Gsl + GG + GM + GA) to alter
the missile's trajectory thereby guiding the missile to the
target with proportional navigation.
\~hen
the missile is carried on the aircraft in a nontracking
mode, the seeker is constantly constrained (cage) to approximately
zero degrees seeker angle.
The caging loop processes the seeker
angle through a switch in the launcher and G52 to produce a
torque on the seeker gyro which reduces the seeker angle to zero.
The GCS actually processes the seeker loop information in
circular coordinates which requires only a single channel for
azimuth and elevation information, and processes the guidance
information in cartisian coordinates requiring two processing
channels, one for azimuth and one for elevation.
The change
from circular to rectangular coordinate processing occurs in GG
by quadriture demodulating the error signal from G51 with the
gyro reference signal.
The caging loop gain (G 52 + G53 ) was measured (with no
target input) by precessing the seeker angle to thirty degrees
off axis and measuring the time required for the seeker angle
to be reduced to ten degrees upon closing of the caging loop.
The method of precessing the seeker was to apply a phase shifted
sample of the gyro reference signal to the input of G52 as shown
with dotted lines in Figure 17.
The guidance loop gains G51 and GG were monitored in both
azimuth and elevation channels by tracking a target with a certain
46
angular rate and monitoring the electrical output of G6 . The torque
motors which provide canard rotation were not energized due to the
fact that a mechanical torque measuring device could not be attached
to the motor outputs without altering the missile•s response to
acoustically induced vibration.
The cagin loop test and guidance loop tests together monitor
the critical system gain G , G , G and G . Additional tests
6
51
52 53
were performed to monitor nonlinear functions in the processing.
The functions and methods for measuring thosefunctions are described
below.
Lambda Compensation is a bias applied to the guidance drive
signal that rotates the heading of the missile into the direction
of the target during the first few seconds after the missile is
launched (i.e., during the time the missile is boost accelerated).
This keeps the missile trajectory from overshooting the target
during the time the missile velocity is changing very rapidly.
Lambda compensation was measured in both axes by tracking a
stationary target at thirty degrees off axis and monitoring the
outputs of GG for the first few seconds after simulating in the GCS
a launch condition.
Rate Bias is a function, nonlinearly proportional to the seeker
angle during free flight, that applies a small torque to bias the
seeker angle toward zero degrees.
This alters the missile trajectory
and causes the seeker to tack the most forward part of the target's
infrared signature.
The rate bias was measured by precessing the
seeker off axis thirty degrees and alters several seconds after
47
a launch condition measuring (1) a telemetered signal from the rate
bias circuitry and (2) the bias at the output of GG.
Torque Limiting is a feature included in GG to provide overall
guidance loop stability in the end game portion of the flight when
very large guidance accelerations are generated to intercept the
target.
The torque limiting was measured in both axes by monitoring
the output of GG with the seeker tracking a target that has angular
movement sufficient to produce an output from GG that is clamped by
the torque limiting circuitry.
Several parameters were measured to determine the condition of
the mechanical gyro (seeker) and the electronic portion of the
GCS which dive the gyro rotation.
The gyro operating speed was
monitored,to indicate proper operation of the speed control circuitry,
by measuring the frequency of the gyro reference signal.
Accuracy
of this frequency is critical in the seeker and guidance loop
processing.
The run-up time measured from application of missile
power from zero rpm to normal operating speed, and run-down time
measured from removal of operating power was measured to give an
indication of gyro bearing condition.
A Dark Cell Noise test measured the noise in the infared
detector section to also give an indication of gyro bearing
smoothness of operation, and more importantly to give an indication
of the target detection sensitivity.
The Dark condition for this
test was approximated by masking all mechanical objects in the
seeker 1 s field of view, including the chamber walls, with lamp black
curtains or paint.
48
A test to measure overall seeker tracking performance was
implemented by measuring the maximum target angular line of sight
rate of motion at which the missile would maintain track.
This
was accomplished by slowly increasing the rate of target motion
until the seeker lost the tracking signal.
The audible signal (chirp) which is presented to the aircraft
pilot to indicate that the missile is tracking, was also monitored
by the test set.
External power which is the normal captive flight
mode of the missile, and internal power which simulates the missile
being powered by an internal battery during free flight, were both
monitored for excessive current drain which would indicate a
malfunction in some part of the GCS.
The combination of these tests checked essentially every
function of the GCS operation (except canard torque motors and
the internal thermal battery) and during the AIM 9L NTE detected
and identified every GCS failure.
Target Simulation
The method of simulating a moving target for the GCS functional
testing was determined by first defining the target requirements for
each test as shown in the table below.
Target Requirements
Functional Test Performed
External and Internal Power
None
Gyro Speed - Regulated, Up, Down
None
Cell Noise - Semi-Dark
No Target - Black Surrounding
Caging Loop Gain
No Target
Maximum Tracking Rate
Variable to 25°/sec on axis
49
Target·Reguirements
Functional Test Performed
Autopilot Gain
Moving XOfsec
Torque Limiting
r~oving
Lambda Compensation
Fixed 30° off axis
Rate Bias
No Target - Black - Surrounding
Chirp
Moving X0 {sec
yo {sec
From this information and knowledge of the seeker operation the
following requirements were defined for the target simulation design:
1.
Target must be capable of being turned off, with dark
surroundings.
2.
Target must move from at least 0° to 30° off axis.
3.
The 300 off axis stationary target must pass through 0°
to acquire.
4.
The target must have variable angular motion from xo;sec
to 25°/sec.
5.
The 25°/sec motion should occur around 0°.
6.
The target must move through both azimuth and elevation
axis of the missile.
In order to satisfy these requirements for target motion two
possibilities as shown in Figure 18 immediately came to mind.
1.
Initiate track of the target at the origin (acquire),
translate the target to some radius and circulate the target around
the origin (zero degrees azimuth and elevation).
2.
Linearly move the target with sinusoidal motion at 45° to
the missile axis through the origin to both acquire and track.
The latter approach was selected and implimented using a
50
-_,.'
.
'
APPROACH #2
k'---r---,
i
'\.
\
\
APPROACH #l
FIGURE 18 - APPROACHES TO
..
SIDEt~INDiR
SIMULATED TARGET MOTION
51
simple motor driven crank assembly shown in Figure 19 to drive a
horizontal metal rod to which the targets were attached.
The GCS's
were rotated 45° to allow the horizontal motion to exercise both
azimuth and elevation channels.
Due to the constrafnts
of
the chamber dimensions, two separate
infrared sources It/ere used with each missile to satisfy a11 of the
target requirements.
One source was mounted on the moving metal
rod to provide target motion from 0° to 30o off axis.
The second
source was mounted to provide plus and minus 15o motion with the
maximum angular rate of motion at the origin.
The test set
energized (with a fixed voltage) one source or the other for each
missile depending on the test being performed.
The test set also controlled the speed of the target motion
motor.
Microswitches mounted to the motor stand were interfaced
to the test set to sense target poistion for halting the target
motion at 30°, and for measuring certain parameters that require
a specific target angle or angualr rate.
The rotating speed of the
motor, which is directly related to the angular target motion, was
measured very accurately by using a toothed disk attached to the
motor shaft to interrupt the light path between a light emmiting
diode and a photo cell, and noting the frequency of interruptions.
The peripheral devices used to measure the tested parameters
were:
1.
Hewlet Packard 3400A True Root-Mean-Square voltmeter to
measure cell noise during the detector sensitivity test.
2.
Darcy model 330 digital multimeter used on the 100 volt
52
r·-------- ----- --------
-----------~
!
1
\ /
.....----+----~ ~ ,.,.--'--:~--
!
FIGURE 19 -
SID~WINDER
SIMULATED TARGET MOTION USED FOR
THERMO-ACOUSTIC RELIABILITY TESTING
I
•.
53
scale to measure
± 11
volt GCS internal power, rate bias, lambda
compensation bias, guidance loop gain and torque limiting.
3. Anadex mbdel CF-200 countertimer used in the external
programmed countermode to measure gyro operating speed, gyro runup time, gyro run-down time, rate bias precession time, caging loop
response time, and target motion angular rate.
4.
Hewlet Packard model 122-AR oscilloscope used to visually
display the azimuth and elevation coordinates of the seeker angle
or the canard torque motor driving signal.
All of the peripheral devices were operated in a single mode
or range and required no adjustment during the functional testing.
In addition to the peripheral measuring devices the GCS
functional test set itself had several measuring devices as seen
in Figure 20.
Four panel meters monitore GCS external and internal
power and an amplifier and loud speaker monitored the audible
(chirp) tracking signal.
Labeled panel lights indicated which peripheral device to
read and the units of measurement for recording on the GCS functional
test log sheet reproduced as Figure 21.
Labeled lights also
indicated whether the missile was externally or internally powered
and whether the seeker was tracking or caged.
Lights were also
used to indicate target motion, which target is energized, and to
indicate the angular target motion rate which should be adjusted
by the operating personnel to either six degrees per second or
increased slowly to maximum tracking rate.
The test set was designed using a minimal number of different
54
FIGURE 20 - SIDEWINDER GUIDANCE AND CONTROL SECTION
FUNCTIONAL TEST SET MAIN FRAME
~
MISSILE #
.,
......
Gl·
c:
;:o
DATE
TEMP
-v
+I
STATION
-v
+llv
Speed
AI~l
Run ll\•m -11V
PAGE
LAMBDA-
CAGF.
DCN
R UP
9L LOG
RATE B TRK
TIHE
CHIRP
f.J..
EL
-v
+V
+I
-I
OF
GAIN
AZ
EL
1.Tral
AZ
EL
STATUS INITIAL
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56
components.
Standard off the. shelf electronic components with
normal tolerances were used whenever possible.
Precision resistors
when necessary were mostly the same value of 100,000 ohms. ·
The main building block of the circuit was a Raytheon RC 4136
quadruple operationaT amplifier in a 14 pin dual-in-line package.
The independent operational amplifiers, similar to the standard 741
were used as:
1.
common mode rejection differential
input buffer/amplifier
.
2.
voltage comparator
3. relay driver
4.
phase inverter
5.
high impedance buffer
6.
voltage translator
7.
line driver
8.
track and hold amplifier
9.
quadriture demodulator
10.
monostable·multivibrator
11.
frequency doubler
The only other integrated circuits used were two Signetics
SE-555 timers and a Signetics uA-741-CV operational amplifier. A
dual-in-line reed relay was used in low power applications and a
six pole double throw general purpose relay, driven by a power
transistor, was used for medium power application.
·oescription·of·Tests
The following is a description of each semiautomatic test
performed by the test set.
The description references the Test
57
Set Wiring and Block Diagram of Figure 22 and identifies in
parentheses one of the four printed circuit cards on which the
function being described is mounted.
The semiautomatic tests are controlled by a 6 pole 14 position
rotary switch.
Due to the fact that the zero position has no
electrical contact with any of the poles this zero position of the
switch was arranged to correspond to the fifth test performed which
required the least control of internal and external functions.
Therefore, test number one corresponds to switch position ten, test
number six corresponds to switch position one, test number fourteen
corresponds to switch position nine.
The counter (far right hand side of Figure 22) is controlled
by the test set to always reset the display on the leading edge
of the enable signal by generating a 1 msec (lo- 3) reset pulse with
a One-shot (2).
Unless otherwise specified the GCS seeker is caged by applying
the caging Voltage through relay Kl (3) to the cagingreturn.
TEST #1
External Power is monitored by a test set voltmeter connected to
the GCS Interface plug pin #30 (far left of figure) and a test
set current meter in series with the +26 volt power connected
through relay KlO to GCS Interface pins #1 and #2.
External power
supplies provide plus and minus 26 volts (rftferenced to ground) to
the test set for powering both the GCS and parts of the test set.
Gyro Speed is monitored as follows:
the U/D Reference signal is
Buffered (3), doubled (2) by full wave rectification, shaped (2) to
58
-·
t II 21
II
~~
~·
13 ~~
'f
.-~
!?C.:.
-~:
H-
t:-
t~DI
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F
~~
s-
I
'i-
,.. cr
3-
6 110 /0
C:iC 5 17
'F-
C-
AUO 10 18
Iii
~
-
AliTo,
u:s
0 7
CARD•Li l
RE6
LO AD I~ r--
3~
f - - 'i
(rl"' ~L
I··
r-2.
.-I
1-
i
i
-
I
I
/7:~ r;·
EX T 2.
.~
'
---,
:..:
HT Z3
NO 2
N
lz.q
bv
&
}
~ i30 f---<0--
6·AI&
-EIr31
-=
v~;-
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k0-
!.
-~
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GZ-5
TER.FACE
6No
'-v-'
...._.
COOL AIR
FIGURE 22 - SIDEWINDER GUIDANCE AND CONTROL SECTION
FUNCTIONAL TEST SET WIRING AND BLOCK
DIAGRAM.
59
provide fast rise and fall times and counted by the external counter
connected to the Count Frequency jack. The counter is reset and
enabled for five seconds by the leading edge of a 0.1 Hz square
wave and is disabled and displays the data for recording by the
operator for five seconds at the trailing edge of the 0.1 Hz square
wave.
At the end of the five second display period the counter is
reset, and the process repeats, etc.
+11 Volt internal power is Buffered (2) and applied to the external
Digital Voltmeter which is connected to the DVM jack.
TEST #2
Run Down Time is measured by removing missile external power with
relay KlO and resetting and enabling the counter to count a 10Hz
signal supplied by the counter, until the frequency of the
Buffered (3} U/D Reference signal is less than 20 Hz as determined
by the Frequency Detector circuitry (2) at which time the counter
is disabled and displays the time in tenths of seconds.
TEST #3
Run Up Time is measured by applying missile external power with
relay KlO and resetting and enabling the counter-timer to count
a 10 Hz signal supplied by the counter until the frequency of the
Buffered (3) U/D Reference signal is ninety five percent of the
operating frequency as determined by the Frequency Detector
circuitry (2} at which time the counter-timer is disabled and
displays the time in tenths of seconds •
..:.11 Volt internal power is Buffered (2) and applied to the Digital
Voltmeter through the·DVM
jack~
60
TEST #4
··caging loop response is
m~asured
as follows:
a logic signal from
card number one Control Logic is applied to the Logic (3)
and Seeker Angle Threshold (3).
This opens relay K91 (3) and
enables a phase shifted (3) sample of the U/D Reference signal
applied to the caging return to precess the seeker gyro in an
up-left direction.
The Buffered (3)caging voltage, whose amplitude
is proportional to seeker angle, as applied to the Seeker Angle
Threshold (3) circuit which determines when the seeker is precessed
to 300 off axis at which time it 1.
resets and enables the counter-
timer to count a 10 KHz signal and, 2.
which causes the seeker to cage.
de-energizes relay Kl (3)
When the seeker angle is 10°
off axis the SeekerAngle Threshold (3) disables the counter which
holds and displays the time in lo- 4 seconds.
The caging voltage whose amplitude is proportional to angle
magnitude (radius in circular coordinates) and whose phase is
related to direction (angle in circular coordinates) is quadrature
demodulated with the U/D Reference signal in the Seeker Position
Demodulator (3) and the azimuth and elevation information displayed
on the horizontal and vertical axis respectively of the external
oscilloscope.
TEST #5
Dark Cell Noise is measured in rotary switch position zero by applying the Buffered (3) GCS Audio signal to the True Root-Mean-Square
voltmeter through the_ TRMS
jack~
Rate Bias measurement is initiated by applying internal missile
61
power through relay Kll and applying the Rate Bias signal from the
GCS Interface to
th~
DVM.
One sec6nd later a Delay circuit (4)
removes external power with relay KlO.
Two seconds later the
control logic (1) applies a signal to the Logic (3) and Seeker
Angle Threshold (3).
This opens relay K31 (3) and enables a
phase shifted (3) sample of the U/D Reference signal applied to
the caging return to precess the seeker gyro in an up-left direction.
The Buffered (3) caging voltage, whose amplitude is proportional
to seeker angle, is applied to the seeker Angle Threshold (3)
circuit which dtermines when the seeker is precesses to 300 off
axis at which time it, 1.
resets and enables the counter-timer
to count a 10 Hz signal and, 2.
deengergizes Kl (3) and, 3.
forces
the Track Detector (3) circuitry to command the missiles to track.
The Rate Bias voltage is directly read on the DVM and the seeker
angle is slowly reduced. by the rate bias.
When the seeker angle
is 150 off axis the Seeker Angle Threshold (3) disables the counter
which holds and displays the time in tenths of seconds.
The
azimuth and elevation seeker angle information is presented to the
oscilloscope as described in the Caging loop response test.
TEST #7
Track and Chirp is initiated by removing internal power and applying
external power (deenergize relays KlO and Kll), removing the
tracking command from the precious test, energizing Target B
(which is positioned on the target drive to move from 0° to 30°
up-left), initiating the target motion and illuminating the
indicator light which specifies the operator to set the target
L
62
motion to 6°/sec.
The target motion rate is adjusted by the Target
Speed control (lower middie of figure}.
The target motion speed
is dispayed by counting the frequency of the Motor Tachometer
(far right of figure) for 0.7 seconds as controlled by a 1.4 Hz
square wave Oscillator (2} which repeatedly resets and enables
the counter for.0.7 seconds and displays the count for 0.7 seconds.
When the target position is at 0° the GCS seeker will produce
a GCS Audio signal that, Buffered (3}, is applied to the Track
Detector (3} which commands the missile to. track the target. The
track command opens a relay in the caging return line just
exter~al
to the missile in a Missile Cable Buffering Unit which prevents
the capacitance loading of the long cables from altering the
characteristics of the electronic processing in the GCS.
When
the seeker is tracking, the Missile Audio signal is Buffered and
Amplified (3} and applied to a test set speaker which allows the
operator to audabally monitor the signal.
When the target is 300 off axis the goo microswitch attached
to the target motion drive assembly signals the Control Logic {1)
to halt the target motion.
for the following test.
This generates the proper conditions
If at any time the seeker looses track
the Track Detector (3} will remove the track command causing the
seeker to cage, and will through the Control Logic (1} initiate
target motion again to repeat the tracking sequence.
TEST # 8
Lambda Compensation·Elevation is initiated by applying internal
power and removing external power with relays KlO and Kll.
Down
63
and Right torque commands from the GCS are Buffered (1) and applied
to the vertical and horizontal inputs of the oscilloscope.
During
the first three seconds after external power is removed lambda
compensation biases the Right (and Down) torque commands as is
noted on the oscilloscope and measured on the Digital Voltmeter.
TEST # 9
Lambda
Compensation·Ati~Oth
this test, which must be completed within
three seconds of initiating Test #7, continues the conditions of
Test #7 but measures Right torque command with the DVM.
If the
test is not completed within the three second time period, the
missile may manually be Externally Powered with the Interna-lExternal-Automatic switch for five seconds and then returned to
the Automatic mode which will again internally power the GCS and
producing Lambda compensation for three seconds.
TEST #10
System Gain Elevation is measured with the GCS internally powered
and is initiated by extinguishing Target B, illuminating Target A
(which is positioned on the target motion drive to provide
± 15°
motion), and initiating target motion at 6°/sec. The seeker
loses track on target B, cages as described in Test #7 and initiates
track on Target A when-it passes through 0°.
When the target
motion is 20/sec the soo microswitch mounted on the target motion
drive through the Control Logic (1) commands the S &H (1)
to Sample and Hold the value of the Buffered (1) Right torque
signal and display this voltage on the DVM.
64
TEST #11
System Gain Azimuth maintains the same.conditions of Test #10 except
rely
K3 (1)
selects the Buffered (1) Down torque signal to be
Sample and Held (1) and displayed on the DVM.
TEST #12
Torque Limit
Eley~tion
maintains the same conditions of Test #10
except the oo microswitch controls the Control Logic (1) to
Sample and Hold and display
~n
the DVM the Buffered Down signal
\)
when the target is moving at 60/sec.
TEST #13
Torque·Limit Aiimuth maintains. the same conditions as Test #12
except the Control Logic (1) commands relay K3 (1) to sample and
hold and display the Buffered (1) Right torque signal when the
target is moving at 6°/sec.
TEST #14
Maximum Tracking Rate is measured by slwoly increasing the target
motion speed until the seeker loses track as is determined by the
Track Detector (3) circuitry which illuminates the Track indicator
light during tracking and extinguishes the indicator light at
loss of track.
The maximum target angular rate of motion is then
noted from the counter as described in Test #7.
A rear view photograph of the GCS test set is shown in
Figure 23, indicating the location of the"printed circuit cards
and the general construction of the unit.
65
FIGURE 23 - SIDEWINDER GUIDANCE AND CONTROL SECTION
FUNCTIONAL TEST SET MAIN FROME REAR VIEW