A new launcher has been developed by DSI, that is compatible with the GAS canisters. It has the proven capability to deploy a satellite from an orbiting Shuttle that is 18 inches in diameter, 31 inches long, and weighing 190 pounds. These DSI Launchers were used aboard the Discovery (STS-39) in May 1991 as part of the Infrared Background Signature Survey (IBSS) to deploy three small satellites known as Chemical Release Observation (CRO) satellites A, B, and C. Because the satellites contained hazardous liquids (MMH, UDMH, and MON-10) and were launched from GAS Cylinders without motorized doors, the launchers were required to pass NASA Shuttle Payload safety and verification requirements. Some of the more interesting components of the design were the V-band retention and separation mechanism, the separation springs, and the launcher electronics which provided a properly inhibited release sequence operated through the Small Payload Accommodations Switch Panel (SPASP) on board the Orbiter. The original plan for this launcher was to use a motorized door. The launcher electronics, therefore has the capability to be modified to accommodate the door, if desired

Gibbons, D.

Nguyen, D.

Nichols, S.

Wise, J.

English

NASA Technical Reports Server

N93-13180
THE DSI SMALL SATELLITE LAUNCHER
S. Nichols, D. Gibbons, J. Wise, D. Nguyen
DEFENSE SYSTEMS, INC.
McLean, VA 22102
ABSTRACT
A new launcher has been developed by DSI, that is compatible with
the GAS canisters. It has the proven capability to deploy a
satellite from an orbiting Shuttle that is 18 inches in diameter,
31 inches long, and weighing 190 pounds.
These DSI Launchers were used aboard the Discovery (STS-39) in May
1991 as part of the Infrared Background Signature Survey (IBSS) to
deploy three small satellites known as Chemical Release Observation
(CRO) satellites A, B, and C. Because the satellites contained
hazardous liquids (MMH, UDMH, and MON-10) and were launched from
GAS Cylinders without motorized doors, the launchers were required
to pass NASA Shuttle Payload safety and verification requirements.
Some of the more interesting components of the design were the V-
band retention and separation mechanism, the separation springs,
and the launcher electronics which provided a properly inhibited
release sequence operated through the Small Payload Accommodations
Switch Panel (SPASP) on board the Orbiter.
The original plan for this launcher was to use a motorized door.
The launcher electronics, therefore has the capability to be
modified to accommodate the door, if desired.
INTRODUCTION
DSI developed the launchers because the presence of the
hazardous payload required more safety inhibits than existing
designs could provide. As the satellite design evolved, more
volume and weight capability was needed. The new launcher has the
electronics and spring drive outside the GAS canister envelope.
This paper will describe the design features, the qualification
program that was performed that led to the approval to fly on the
Shuttle, and the operational results.
The DSI launchers were used aboard the Shuttle, Discovery
(STS-39) in May 1991 as part of the Infrared Background Signature
Survey (IBSS). Fig. 1 shows the canisters and launchers mounted in
the Shuttle payload bay. The SPASP was mounted on the aft flight
deck, as seen in Fig. 2.
These launchers can be reused. Information is provided as to the
processes required for reuse and possible modifications to adapt to
other requirements.
221
PR_C.,,EDING PAGE BLANK r_O'F FILMED
https://ntrs.nasa.gov/search.jsp?R=19930003992 2020-03-17T10:13:47+00:00Z
DESIGN
Mechanical Details
The DSI launcher was designed to interface directly to the
standard GAS canister cylinder. The launcher base plate interfaces
to the cylinder via the standard attachment of 32 3/8" screws. The
mounting to the Shuttle is a typiCal GAS canister technique, using
4 attach brackets and a shear pin.
The primary components 0f the launcher mechanism, as Sh0wn in
Fig. 3A, are the pusher plate, v-band, moun£ing ring, separation
springs (2), base plate, guide rod, retention springs,
electronics/battery boxes, and the electronics enclosure. Most of
the parts are made of aluminum. The springs are made of music wire;
the guide rod is stainless steel; the bushings supporting the
guide rod are bronze.
The launcher pusher plate sits inside the launcher mounting
ring and mates with the inside edge of the satellite mounting ring.
The segmented v-band, as shown in Fig. 3B, holds the rings together
with approximately 3000 ibs tension. The springs were sized to
provide a nominal separation velocity of 3.7 ft/sec for a 190 Ib
satellite. With these springs, the separation velocity will vary
with the satellite weight. All components and hardware are
mounted directly to the launcher base plate.
The launcher electronics system is contained in a single box
as shown in Fig. 4. The batteries and electronics are mounted
beneath the base plate in the payload bay, and protected by the
electronics module cover plate, which provides an additional 8
inches of usable height inside the GAS can, while utilizing only 2
inches of height inside the canister for launcher volume.
The DSI launcher utilizes a 4-segment, aluminum v-band, using
A-286 CRES links at two joints and separation pyrotechnics at the
other two joints. The launcher mechanism is released by firing
clamp separators, actuated by NASA Standard Initiators (NSI). Each
half of the separated v-band is retained in the cylinder after
deployment by retention springs. Damage to the GAS cylinder is
prohibited by band stops, which retain the v-band and prevent
flipping and resulting contact with the cylinder walls. As the v-
band separates from the satellite and launcher mounting rings, the
spring/pusher plate then pushes the satellite out of the cylinder.
Fig. 5 is a photo of a launcher in flight configuration, after
deployment.
Electrical Description
The launcher electronics system accommodates thr@e i_unchers
with control by the SPASP. As shown in Fig. 2, this panel has 6
switches and 6 indicators. Two of the switches were used for the
launcher address code, three were used for the remaining three
inhibits, and the final switch, "execute", was used to initiate the
command selected by the other switches. The indicators were used
to verify the commands.
222
The launcher electronics block diagram is shown in Fig. 6.
Two 12-volt batteries provide power, each with six lead-acid cells
with 4 amp-hour rating. Power converters include a 12V-28V unit for
the indicators on the SPASP and a 12V-SV unit for the logic and
relay drivers. _he pyro circuit uses the 12 volts from the
battery. Both batteries and both converters are protected by
fuses.
The logic circuit is a set of static gates that convert switch
actions into responses to satisfy operational and safety
requirements. The circuit design does not allow the failure of a
single device to cause an improper action, such as a premature or
simultaneous release.
Toggling the execute switch turns the electronics on for 15
minutes. The logic circuit decodes the address switches and closes
the select relay if the address is correct. The prearm switch
closes the prearm relay which is on the return side of the pyro
circuit, and a one minute inhibit timer is started. This timer
prevents an premature release by inadvertently throwing the
switches. The arm switch closes the arm relay. The deploy switch
completes the pyro circuit and releases the satellite from the
launcher.
Ground Support Equipment
Pallets were developed to support the CRO satellites and the
DSI launchers. They were used during the fuel and oxidizer loading
operations on CRO, for transport of the satellites, during the
mating of the satellites to the launchers and the installation of
the V-bands, and through final flight preparation until integration
into the GAS cylinders.
The battery charger provided monitoring and charging of the
batteries during testing, integration and prelaunch activities.
Final battery charge is Performed up to 2 months before launch.
The ordnance circuit monitor unit (OMU) is connected to the
launcher in place of the Safe or Arm plug. This connection safes
the ordnance by shorting the wires while monitoring the activity on
the launcher ordnance circuit. When the fire voltage is applied,
a light comes on in the OMU. This unit also verifies no operation
during the environmental tests.
The function and power monitoring unit (FPMU) is used in
conjunction with a SPASP to verify operation of the four inhibit
relays and to monitor battery and converter voltages. Proper timer
operation and addressing functions may also be verified using the
FPMU.
An interface verification test (IVT) set was developed to
verify the copper path of the interconnect wiring system from the
SPASP to all three launchers. Each launcher has two connectors for
system wiring which are wired together, with branch wiring to the
launcher electronics. The first two launchers each have two cables
connected, while the third launcher only has one cable connected.
223
By attaching the IVT set to the unused connector on the third
launcher, all wiring could be verified in one test without
disturbing flight connections. A test (GSE) SPASP was built to
allow operational tests before integration with the Shuttle.
SHUTTLE QUALIFiCATiON PROGRAM
Design Verification Tests
These tests verified that the operations were in compliance
with predicted and required performance characteristics. The GSE
SPASP was used with two flight cables and one non-flight cable.
All credible scenarios were verified.
A low voltage test was performed to verify the ability of the
logic to perform as the supply voltage decayed. The test result
showed that the relay drivers ceased operations at a higher voltage
than the logic.
Envlronmental Tests
EMI and EMC testing was performed to verify the ability to
operate in the Shuttle RF environment and that no inadvertent
emissions from the launcher could effect the Shuttle.
The Thermal Cycling Test was performed to verify operation of
the launchers over the temperature range of -30°C to +40°C. The
Thermal Vacuum Test, performed at GSFC, was a 90 hour duration test
with holds of up to 8 hours at both -20°C and +40"C, the low and
high temperatures. The Vibration Test, was conducted with the
launchers in the GAS cylinder with the CRO satellite attached in
flight configuration. The vibration levels were 5.5 Grms,
7.7 Grms, 7.0 Grms in the X, Y, & Z axes respectively, as specified
in the Orbiter Cargo Bay Random Vibration Payload Sidewall
Adapters/Orbiter Interface Per ICD-A-17559. The Static Load Test,
performed at GSFC, tested the launcher structure and separation
mechanism, and showed minimum safety factors of 2.78 on yield and
3.34 on ultimate.
Operatlonal Verification
As part of the design process, an extensive deployment
velocity analysis was performed. The goal of the analysis was to
provide a nominal separation velocity of 3.7 ft/sec for a 190 ib
satellite. Deployment testing was performed to verify the
analysis, and some surprising results were found. Deployment
testing showed a separation velocity of approximately 60% the
predicted value. The losses were discovered to be due to dynamic
response of the spring mass system. All previous analysis
considered only static properties. Design modifications were
incorporated to achieve the required deployment velocity.
An analysis was required to ensure the launcher would not
"jam", creating an unsecured, undeployed sateiiite Situation. The
inherent design prevents the launcher from jamming itself.
Analysis also showed that all possible angles of incidence of the
F
224
satellites and the cylinder walls, given the coefficients of
friction (minimized by placing delrin rails on the outer edges of
the satellite) were too small to cause the satellites to jam within
the cylinder after separation from the launcher. Flight experience
shows that all three CRO satellites contacted the cylinder wall
with no jamming.
Tipoff analysis was required to determine expected tumbling,
interference, etc. The GAS cylinder walls limit tumbling due to
tip-off, but is still possible, as post flight data indicates.
Tip-off is a function of satellite mass properties, so a tip-off
analysis must be performed for each mission. Post flight data does
not correlate to previous analysis well, simply because tumbling
was not of concern to the first mission.
V-band release testing was performed in numerous
configurations to verify the structural design, the release
capability, and the retention test. Ground tests verified that the
v-band would release by firing only one of the two clamp separation
bolts. Ground testing also verified the v-band attachment
procedures as well as leading to the final v-band retention design,
which was difficult to design based on analysis alone. The v-band
was also test fired using bolts and a bolt cutter, instead of the
recommended clamp separator. Tests proved the band released
nominally in this configuration.
Flight Safety Qualification
The DSI launcher has passed extensive flight safety
qualification and certification in accordance with NSTS 1700.7B,
the STS Flight Safety Handbook. Extensive electrical, structural
and kinematic analyses and tests were performed to certify that the
launchers would pose no operational threat to the Orbiter. In
addition, the launcher handling is controlled by a certified and
approved Fracture Control Plan, generated in accordance with NHB
8070.1, the STS Fracture Control Requirements. In the event of a
re-flight, the Fracture Control Plan is still applicable for all
handling and screening of flight hardware.
OPERATIONAL RESULTS
On-Orblt results
CRO satellite C was successfully deployed with a confirmed
ejection velocity of 4.0 ft/sec which was within 5% of the
predicted release velocity of 3.9 ft/sec. No operational problems
were encountered during the deployment sequence.
CRO satellite B was successfully deployed as shown in Fig. 7.
The measured release of 3.7 ft./sec, was within 5% of the predicted
release velocity of 3.6 ft/sec. Upon initial attempt to deploy B,
the astronauts were unable to activate the launcher B electronics.
After following the malfunction procedures, there was still no
response from the launcher. Groundcontrol commanded the crew to
try launcher B again, before going to launcher/satellite AM (which
was the planned contingent mode). Launcher B performed nominally
225
and satellite B was successfully deployed.
discussed in the next section.
This anomaly is
CRO satellite A was successfully deployed, with a measured
velocity of 3.5 ft/sec, which is within 5% of the predicted release
velocity of 3.4 ft/sec. Problems similar to those seen in the B
deployment were observed here. Again, this anomaly is discussed
in the next section.
Post Flight Aotivities
When the Shuttle returned to the Kennedy Space Center (KSC),
a series of tests were performed during the deintegration as part
of the anomaly resolution. In the Orbiter Processing Facility,
operational and cable resistance tests were performed before any
changes were made to the flight configuration. (Repeating the
verification tests performed pre-flight.) Both tests were
successful and showed no anomalies.
Stand-alone tests were performed by DSI both at KSC and in
McLean, VA using the GSE. All tests were successful and no
anomalies were found. The Shuttle payload bay operated at a cold
temperature during the flight of STS-39. Therefore a cold
temperature test was performed to simulate flight conditions.
This test recreated the flight anomaly.
The flight anomaly was caused by the design of the electronic
circuit. The execute switch is used for two functions: To turn on
the electronics for 15 minutes; and to initiate switch selections.
Due to a time delay caused by the toggling of a relay, the input
voltage to the logic is applied before the logic is powered on.
When the temperature is cold, this caused the logic to "lock-up"
and not respond to further commands. Resetting the logic clears
this condition and allows the logic to perform normally. The
proper operation, in accordance with the electronics design, is to
toggle the execute switch by itself to turn the launcher
electronics on, and then to set the address switches and toggle the
execute switch again.
Future Mission Capabillt7
Refurbishment Requirements
Because the DSI launcher is flight proven, a minimal amount of
refurbishment is required. In order for a payload to use the DSI
launcher, there is a certain minimum hardware refurbishment and
engineering support required.
Hardware Activities:
- Battery replacement.
- Mounting ring & v-band inspection and surface re-finish.
- Procurement of new clamp separators (although only 2 are
required for flight, at least 20 must be procured due to STS
ordnance qualification requirements).
- Standard testing (functional and temperature).
- Separation testing.
226
F
Analysis, Documentation, and Program Support:
- Separation analysis, which is spacecraft unique.
- Flight Safety Documentation. According to NSTS 1700.7B, must
be re-certified for each flight. All analyses for flight 1
are in place, so it is a matter of verifying analyses are
still adequate, re-submitting documentation as part of payload
unique safety process.
- Support of system vibration testing and pyro shock testing (if
customer required).
- Support Orbiter integration at KSC.
- Provide crew training (for SPASP operation).
- Support mission operations at JSC if required.
Baseline Capability
The baseline capability of the DSI launcher is a 3.7 ft/sec
deployment velocity for a 190 Ib payload. The separation velocity
vs payload weight, using the standard separation springs, is shown
in Figure 8. If the deployment velocity is too great, smaller,
standard springs may be substituted for the existing ones. If
additional deployment velocity is required, custom springs would
have to be developed.
In addition to deployment velocity limitations, the launcher does
have payload weight vs cg limitations, as shown in Figure 8. These
are based more on Flight 1 qualification levels than structural
capability. If payloads meet this requirement, then little or no
additional structural analysis would be required. If the payload
does not fit the weight/cg curve, additional analysis would be
required.
Motorized Door Option
DSI originally planned to have the capability to operate the
GAS fully deployable motorized door. For this reason, two sets of
high energy batteries were installed and a switch function was
intended to be used to control the door. The DSI launcher could be
used with the door by modifying the electronics. The changes made
would be to add a printed circuit card which would utilize the Pre-
Arm switch as the motor control and the Pre-Arm indicator as the
motor control indicator. The-one minute timer would now be
replaced by the time to operate the door.
Standard GAS Can Launcher
DSI originally intended to utilize the standard GAS Can
launcher (known also as the TSI launcher) marman (V) band design.
If a payload would like to be (or is) compatible with the TSI V-
band, and is in the acceptable weight vs cg range, TSI compatible
separation system could be incorporated at a minimal cost.
227
!FIG I. DSI Launchers Installed in Orbiter DISCOVERY
FIG 2. SPA Switch Panel in Aft
228
Fligh_ Deck of Orbiter
OP,,]GINAL PAGE
BLACK AND WHITE PHOTOGRAPH
m
IF](.] 3. D_][ Launcher Assembty With V- Band
(A) Side View (B) V-Band
/ ° ° °I
o C
[ i
/
FIG 4. DSI Launcher Electronics Box
229
FIG 5. DSI Launcher in Flight Configuration
I EXECUTE SN_TCH t
_561
I FUNCTION C;H_TCHES t
(S!-S_)
1P
_n aus
I ATTERY #2
CONVEI:I!E_
._ TZMZNG CCPlCU_T _J
100_0Nv--TE_--?.l I°_L'Y°°_v_°_
l=:,_o_] r_o._i ,
,ill
II
FIG 6. Electronics Block Diagram
230
FIG 7.
Deployment Velocity i
AllowabLe CG Height
(ft/$oc & inches)
CRO Deployment from Shuttle
3O
2O
10
%
0 50 I00 150 200
Payioad Nass (tlm)
,-';_<:IN_L [)AGE
BLACK AND WHITE PHOTOGRAPH
Deptoyment Vetoci ty
(ft/sec)
.... Haximum AllowabLe CG
FIG 8. DSI Launcher Predicted Deployment Velocities and
Allowable CG Heights as a Function of Payload Mass
231
27L


The DSI small satellite launcher

Abstract

Similar works

Full text

Available Versions

NASA Technical Reports Server