NASA probes damage to fuel tank

Not again!

http://www.cnn.com/2006/TECH/space/0...eut/index.html

Thursday, March 30, 2006; Posted: 10:19 a.m. EST (15:19 GMT)

CAPE CANAVERAL, Florida (Reuters) -- NASA is investigating another mishap
at the Kennedy Space Center, this time an accident involving the remodeled
fuel tank to be used for the next shuttle mission, the agency said on
Wednesday.
Technicians were replacing a vent valve near the top of the 154-foot
(47-meter) tall tank on Tuesday when a Halogen work lamp fell and hit the
tank's foam insulation.

Preliminary inspections show the impact left five small indentations, with
the largest about the size of a stick of gum, and one 6-inch
(15-centimeter) to 7-inch (17-centimeter) long scratch, said Marion LaNasa,
spokesman for tank manufacturer Lockheed Martin Corp.

A detailed inspection of the area was under way, LaNasa said, but the
incident was not expected to affect the shuttle's targeted July 1 liftoff.

The affected area is not among the sections of tank foam insulation that
were redesigned after the 2003 Columbia disaster and again after the July
2005 flight of Discovery, the only launch since the accident.

A piece of foam insulation that fell off the tank and hit Columbia's wing
during liftoff was responsible for heat shield damage that led to the
ship's destruction and the loss of seven crewmembers during atmospheric
re-entry on February 1, 2003.

A similar problem occurred during Discovery's liftoff 2 1/2 years later,
though the shuttle escaped damage.

NASA is preparing Discovery for launch again, but it first must prove that
the new tank design is safe to fly. A series of wind tunnel tests and
analyses are under way.

Safety has been a top priority for NASA, particularly at the shuttle
processing center in Florida, where a series of mishaps have resulted in a
death, equipment damage and several near-disasters over the past month.

Listed below are 2 examples of the caibs lack of correlation of actual
data detected on coumbias ascent jan16, 2003, and foam impact damaging
columbias wing to the extent claimed in their foam impact theory.

#1 Please see graphs pertaining to the accelerometer on the left wing
elevon (V08D9729A), (Caib report vol5 part d13 page 604,605)
At 0.13 seconds after impact excitation the magnitude of the wave
detected by the accelarometer from impact testing is matching the wing
3rd bending mode in magntude and period, where the signal should have
been detected by columbias on board accelerometers on jan 16 2003.
Please note foam impact FI=0 at MET +81.9 seconcd, and the graph listed
by the caib on page, shows the wave closest matching the wing 3rd
bending mode and dampening at a slow decreasing rate, meaning at MET
+82 seconds the impact excitation wave was still detectable by
accelerometer (V08D9729A) sampling at 10 times a second. Instead the
caib states “the most compared� is the 2nd wing bending moment,
which completely disregards the closer match of 3rd wing bending moment
to the excitation pattern from the force of impact. Then the caib
states the differences between test results and actual sts-107 flight
data could be caused by difference of impact location, as the reason
for not aknolegding the 2nd wing bending mode.

#2 Foam impact was determined to have occurred at MET +81.9, the
detailed examination by boeing determined excitation in the wing
started at MET +81.7, 0.2 seconds before foam impact, thefefore foam
impact was not the stimulation for the force revealed from boeings’
analysis from data during sts-107’s ascent on jan 16, 2003.

The caib did not provide or demonstrate a correlation to the
columbia’s on board sensors detecting a foam impacting to the
magnitude required for rcc failure on jan 16 2003, and their own foam
impact testing, therefore the caibs theory has not been validated.

caib report vol v part 13 page 608 par 1
“7.3.3.1 Ascent
All STS-107 MADS PCM strain data, without exception, was nominal during
the ascent flight regime. Nosignificant anomalies were noted.
Comparison of ascent strain gage load indicators showed STS-107 ascent
loads to be within the family of previous OV-102 flight experience.
There was no discernable evidence of an impact load to the vehicle near
MET +81.7 seconds. At the PCM sample rate of 10 samples per second, no
such evidence is expected to be present. Both the extremely short
duration of the impact load (0.003 to 0.005 seconds), and the range of
wing modes (6 Hz and above) preclude such evidence. An interesting
signature near this time was evident in some strain gages. The response
was noted on left wing, right wing, and vertical tail gages. Further
study and scrutiny showed that the signature was inconsistent with
impact loading, and attributable to a nominal ascent load response. A
review of accelerometer data did show signatures consistent with impact
loading. This assessment is discussed in Section 7.4.�

Caib report vol 1 page 34 col 2 par 3
“Debris Strike
Post-launch photographic analysis showed that one large piece and at
least two smaller pieces of insulating foam separated from the External
Tank left bipod (–Y) ramp area at 81.7 seconds after launch. Later
analysis showed that the larger piece struck Columbia on the underside
of the left wing, around Reinforced Carbon-Carbon (RCC) panels 5
through 9, at 81.9 seconds after launch (see Figure 2.3-2).�

caib report vol 2 part d19 page 567-568 col 2 par
The wideband FDM data, which because of its more com-plex encoding took
longer to extract from the OEX recorder tape, also showed some
signatures which are indicative of a debris strike near to 82 sec MET.
One of the accelerometers on the left wing elevons, V08D9729A, showed a
single cycle sinusoidal pulse at 81.9 sec MET that was approximately
2 g in amplitude, as compared to a background vibration level which
generally stayed well below 1 g. This is a fairly sig-nificant pulse
which could easily represent a strike of foam debris upon ascent. The
timing and amplitude of this pulse were taken from a preliminary
assessment of the wideband
FDM data that was printed out on a strip chart recorder by NASA at JSC.

Boeing of Huntington Beach performed a more thorough analysis of the
remainder of the wideband FDM ascent data and in general did not find
much that was anomalous. They found that the overall noise levels and
power spectral density (PSD) matched very closely to the data from the
pre-vious flight, STS-109. They noticed that at approximately 40 sec
MET, the vertical stabilizer had some of its higher order modes growing
slightly larger than normal, and this was attributed to some wind
buffeting that was thought to occur around this time. These modes then
decayed shortly thereafter, indicating that the so-called flutter
instability was not becoming excited, as can occur when the wing
bending modes and the fuselage vertical modes coalesce into a single
coupled oscillation. Boeing.s analysis also pointed out that the
recorded accelerations along the longeron were normal. Detailed
analysis of the wideband FDM data over the time frame around 80-85 sec
MET was performed. For the left outboard elevon accelerometer,
V08D9729A, several wing and elevon oscillation modes were found to be
excited during this time, with the strongest being a second order wing
bending mode that matched best to the fundamental component of the
single cycle sinusoidal pulse at 81.9 sec MET. Boeing.s more detailed
time scale showed the period of the single sinusoidal pulse to extend
over 81.70 to 81.74 sec MET, reaching +3.0 g on the positive peak at
81.71 sec MET, and ï€*2.6 g on the negative peak at 81.72 sec MET.
In addition, another accelerometer on the right wing, V08D9766A,
showed a 1.5 cycle sinusoidal pulse response at a slightly earlier time
of around 80 sec MET. This accel-erometer was located at the
coordinates (X1367.0, Y+312.0, Z) towards the middle of the right wing
and was sensitive to Z-axis motion. This accelerometer recorded an
anomalous pulse beginning at 80.22 sec MET, growing to a first positive
peak of +1.5 g at 80.23 sec MET, reaching a negative peak of ï€*1.9 g
at 80.24 sec MET, then another positive peak of +2.0 g at 80.26 sec
MET, before dying away beyond 80.27 sec MET. The best fit to these
peaks was a combination of outboard elevon torsion and the first wing
bending mode. There have not been any explanations offered for the
cause of this right wing accelerometer response. “


Caib report vol5 part d13 page 602
7.4.1.1 Evaluation of Peak Response at ~82 Seconds
An in-depth study was made to investigate if the peak responses
observed at the left outboard elevon accelerometer at ~82 seconds is
due to the debris impact. Normally, sharp spikes in acceleration are
observed at times during the ascent phase of the flight due to
buffeting event(s). The buffeting load is most significant during the
transonic region. However, it still exists at higher Mach numbers,
which resultsin structural excitation. Shown in Figure 7.4-7 is the
left and right outboard elevon comparison for the 10 second period near
82 seconds. The peak response is noticeable only for the left outboard
location. Filtered responses presented in Figure 7.4-8 verify several
wing/elevons were excited at 82 seconds. The 2nd wing bending response
constitutes the largest component of the peak amplitude. In addition,
the responses of 3rd wing bending and elevon torsion modes contributed
to the peak response.

Caib report vol5 part d13 page 603
“To determine if the debris impact can cause the type of responses
observed in the flight data, analyses were performed using the FEM
model of the wing combined with the reduced model of the Orbiter, which
provides the back-up structure’s stiffness and mass (Figure 7.4-9).
An impulse of 3,000 lbs force (with 0.005 second duration) in
Z-direction was applied to the node closest to the RCC panel #8. The
impulse of this magnitude is reasonable for a 1.5 lb object with a
velocity of 530 MPH impacting the surface at 15 degrees inclination.�

Caib report vol5 part d13 page 604
“Shown in Figure 7.4-10 is the recovered acceleration at the left
outboard elevon location from the transient analysis. The FFT (Figure
7.4-11) of the response indicates excitation of several wing modes,
including wing’s 2nd and 3rd bending modes. The filtered responses
shown in Figure 7.4-12 illustrate the 3rd wing bending mode constitutes
the majority of the peak amplitude, while the 2nd wing bending and
elevon torsion modes also contribute to the peak response. The
acceleration computed using the FEM model is shown along with the
flight measured data in Figure 7.4-13. The shapes of acceleration
signatures are comparable at the onset of debris impact. The frequency
from the analysis is higher, since the 3rd wing bending mode is excited
the most compared with the 2nd wing bending mode experienced during
STS-107. More pronounced 3rd wing bending response from analysis could
be attributed to possible deviations from the assumed location and
duration of impact event, and some uncertainty in the FEM models for
higher order wing modes. Nevertheless, similar acceleration signature
and the excitation of higher order wing modes from the analysis
indicate that the debris impact quite possibly could have caused the
peak acceleration on the left outboard elevon at ~82 seconds in
addition to other aerodynamic disturbances, such as buffeting and
shocks. An absence of additional sensors on the left wing make it
difficult to make conclusive remarks.�

This is from the news piece you posted, which directly states the caibs
theory, so yes my reply is rellevant. The point of my post is to
demonstrate a possiblity there is another cause for the loss of sts-107
crew and columbia simply because the caib did not validate it's own
theory.

Thursday, March 30, 2006; Posted: 10:19 a.m. EST (15:19 GMT)
CAPE CANAVERAL, Florida (Reuters)
"...A piece of foam insulation that fell off the tank and hit
Columbia's wing
during liftoff was responsible for heat shield damage that led to the
ship's destruction and the loss of seven crewmembers during atmospheric

re-entry on February 1, 2003."

The caib provides only a basic correlation of hole size from it's
vast testing to the actual flight data. But the caib does not provide
consistency with possible hole sizes (4", 6", 10", 17" ) amongst its
testing or, to the flight day 2 object of or a 140sqin, or to the
actual sts-107 flight data, and therefore does not validate it's own
theory.

Caib report vol 1, page 63 paragraph 2:
"After exhaustive radar cross-section analysis and testing, coupled
with bal-listic analysis of the objects orbital decay, only a fragment
of RCC panel would match the UHF radar cross-section and ballistic
coefficients observed by the Space Surveil-lance network. Such an RCC
panel fragment must be ap-proximately 140 square inches or greater in
area to meet the observed radar cross-section characteristics."

Caib report vol v part 13 page 92 par 6
"4.4.2 Damage Progression Theory and Supporting Aero
Based on the damage assessment and timeline period correlations covered
in Section 4.4.1, the
following is a postulated damage progression theory based on the
results of the aerodynamic
investigation. This damage progression, approached from an aerodynamic
perspective, is
consistent with the working scenario and attempts to maintain
consistency with other data from the
investigation. References are made to figures which include a
combination of aerodynamic
extraction results and the major timeline events noted.
An initial WLE breach (small hole or slot) in an RCC panel exists at
entry interface. By EI + 300 sec
thermal events are occurring internal to the WLE cavity, however no
identifiable aerodynamic
increments are observed."

Caib report vol v part 13 page 521 par 7
"A comparison of the times at which these critical events occur
during the entry is shown in Table 6-7. As expected, failure times are
accelerated for the 10 inch case compared with the 6 inch due to the
higher levels of internal heating. Thermal response of instrumentation
within the left wing of STS-107 have suggested the initial breach
through the spar occurred at 491 seconds after entry interface. With a
predicted spar breach time of 470 seconds, the 6 inch provides a better
comparison to flight data than the 10 inch case. As shown in Figure
6-82, better agreement for the 6 inch damage case can also be seen by
comparing the temperature response of V09T9895 (panel 9 spar rear
facesheet thermocouple)"


Caib report vol v part 13 page 92 par 6
Holes Through Wing
Limited parametric studies of simulated damage in the form of a wing
breach from the windward surface to the leeward surface were attempted
in this facility and were primarily associated with aerodynamic testing
(see Section 4.3.1). Initially, circular holes dimensionally consistent
with the width of a carrier panel (approximately 4 inches full scale)
were placed at the interfaces for carrier panels 5, 9, 12, and 16. The
holes were found to force boundary layer transition on the windward
surface to the damage site. The model and IR setup for the aerodynamic
tests at this point in time precluded imaging the side fuselage. Since
the model also incorporated damage in the form of missing RCC panel 6,
it is believed that effects (if any) from the carrier panel holes would
have been dominated by the disturbance from the missing RCC panel. TPS
damage in the form of a much larger breach through the wing was
attempted, but the side fuselage heating measurements were considered
qualitative due to compromised phosphor coatings on the models that
were used. The holes were orientated normal to the wing chord and were
located near the left main landing gear door. One hole location was
approximately located at the center of the forward bulkhead
(X=1040-inches in Orbiter coordinates) and the second location was near
the center of the outboard bulkhead (Y=167-inches in Orbiter
coordinates). At each location, the wing hole diameter was
systematically changed from 0.0625 to 0.125 and 0.25-inch at wind
tunnel scale (8.3, 16.7, and 33.3-inchfull scale). While the
compromised phosphor coating considerably degraded the image quality,
it was evident that no change in side surface heating was apparent for
any tested combination of location or diameter."

Surface heating on reentry was indifferent to hole size, not plasma
flow internal to the wing. The caib did not resolve inconsistencies,
nor provide correlation and therefore their theory of foam impact
causing a breach has not be proven or validated.

The caib only established consistency with data external to the
vehicle, but when it came to realistically modeling the thermal values
of breach penetrating plasma approx. 2000 deg f, to sensor detecting
temperatures increasing from 80 deg f, to 180 deg f after minutes of
exposure, the caib was not consistent all. The caib's explanation
for the re-entry temperatures located in the main landing gear box
required phrases such as "However, one could argue that this
convective energy then replaces radiative energy but CFD would have to
confirm this" (see below). The physical properties associated with
the caibs theoretical hot air rushing through the wing during
Columbia's reentry in not based on solid science, as they could not
correlate actual detected temperatures to theoretical effects of
alleged rcc damage.

Caib report vol2 part d19 page 571 col 1 par 2
"REMAINING, UNEXPLAINED INCONSISTENCIES
.... This pathway for the hot gas does indeed exist, but the reason for
the gas to take this tortuous path over other directions is not clear,
nor is it understood why the heating effects would be registered by
only a few sensors on the rear wall of the wheel well and not by others
of a similar type and mounting located only inches away."

Caib report volume 5, part 13 page 512,513
"6.3 Wheel Well Thermal Analysis
A thermal analysis was used to compare predicted heating to the flight
data instrumentation summarized
in Table 6-1 for the wheel well. Here a hot gas plume originating from
the wing leading edge spar is
assumed to impact the outboard wheel well wall. The hot outboard wall
then conducts heat into the
adjoining walls and radiates into the main landing gear (MLG) wheel
well and the associated sensors
within.
Table 6-1 - Wheel Well Sensor Summary
MEAS. NO. DESCRIPTION
V58T0125A SYS 1 LMG UPLK ACT UNLK LN
V58T0405A L H MLG STRUT ACTUATOR
V58T0841A SYS 2 L AFT BK SW VLV RTN
V58T0842A SYS 3 L FWD BK SW VLV RTN
V58T1700A L MLG BRK HTR LN 1 SYS 1&3
V58T1701A L MLG BRK HTR LN 3 SYS 2&3
V58T1702A L MLG BRK HTR LN 2 SYS 1&3
V58T1703A L MLG BRK HTR LN 4 SYS 2&3
NSTS-37398 AeroAerothermalThermalStructuresTeamFinalReport.pd f
NSTS-37398 AeroAerot
A thermal math model shown in Figure 6-16 was developed directly from
computer aided design (CAD) models and used to predict the sensor
responses in this presumed scenario. Shell elements were used where
possible for simplicity with the remaining geometry represented by
solid tetrahedral elements. Main Landing Gear (MLG) components were
then thermally connected by combining nodes at joint locations. This
allowed for faster analysis since arbitrarily low conductors (which is
a common method to join components) can significantly reduce the time
step in order to maintain a required accuracy. Internal radiation was
also modeled using the Monte Carlo technique with 16000 rays per node.
All of the nodes representing sensor locations had an initial
temperature corresponding to the flight data. The rest of the
components and the wheel well walls had an assumed initial temperature
of 80°F. The predicted plume heating distribution model is described
in section 5.3.3. The heating was calculated for a 5 inch diameter hole
in the wing leading edge spar assumed to appear instantaneously at
EI+488sec (13:52:17 GMT). The plume impinged upon the outboard wheel
well wall at location xo=1105, zo=309 and at a distance of 56 inch from
the spar. Correction factors were applied for the 31.5 degree off
normal impact angle as well as the internal wing pressure. The center
of the plume had a heating rate of 22.1 BTU/ft2-sec with the heating
dropping off radially from the centerline. Melting of the outboard wall
was not modeled, therefore, once a node reached melting temperature
(935°F) it was then held at this temperature.
The thermal math model was solved using the Systems Improved Numerical
Differencing Analyzer
(SINDA). Predicted temperatures for the hydraulic lines and the strut
actuator were obtained and are
shown in Figure 6-17 through Figure 6-24. Initially, all the sensors
begin to trend towards 80°F from
radiation exchange with the 80°F surrounding structure. At 488 sec
(GMT 13:52:17) the plume heating is applied to the outboard wall and
the temperatures begin to trend upward at a rate dependant on their
view factor to the outboard wall.
Table 6-2 - Sensor location and view factor summary
Sensor Location View Factor
V58T1700A Bottom of strut Good
V58T1701A Bottom of strut Good
V58T1702A On inboard wall Partial
V58T1703A On inboard wall under debris shield Poor
V58T0841A On inboard wall under debris shield Poor
V58T0842A On inboard wall Poor
V58T0125A On upper wall behind structure Poor
V58T0405A Aft inboard corner of wheel well Partial
At first glance sensors V58T1700A and V58T1701 correlate very well with
the flight data. However, past EI+510 sec (GMT 13:52:39) the low mass
honeycomb access panel has reached its melting temperature in the
analysis as shown in Figure 6-25. This would allow hot gases to enter
the wheel well and deposit energy through convection directly on the
MLG components. However, one could argue that this convective energy
then replaces radiative energy but CFD would have to confirm this.
After EI+586 sec (GMT 13:53:55) the wall at the center of the plume
reached its melting temperature as shown in . At this point the area
available for hot gases to enter the wheel well increases rapidly as
more of the outboard wall melts. The assumption of holding the wheel
well wall at its melting temperature (935°F) is no longer valid for
this analysis and CFD is required to determine the hot gas flow inside
the wheel well in order to account for the convective heating. From
this analysis, it is possible to conclude that a portion of the
hydraulic line temperature increase seen in the flight data can be
attributed to radiation from the outboard wheel well wall being heating
by a plume due to a spar breach."


Given that temperature is a measurement of a systems state, the air
molecules could be assumed to start at the same temperatures prior to
anomalous heating at 80 f def in the main lading gear whell well. This
would mean if the caib were correct, a heat source of at least 935 f
deg heated this area at 80 f deg, for a period of more than 6 minutes
and a temperature sensor approx. 60 inches away only registered
increases of approx. 100 f deg.

Below I have correlated hydrualicy fluid temperatures, and the
anomalous temperatures detected in Columbias left wheel well during
reentry feb1, 2003, therefore showing the plausibility of hydraulic
system breach.


EI +923 LOS (begin data reconstruction)
brake line temp sensor "A" L MLG BRK HTR LN 1 SYS 1&3 (V58T1700A)
final temp at 172 f deg.

Brake line temp sensor "B" L MLG BRK HTR LN 3 SYS 2&3 (V58T1701A)
final temp at 154 f deg.

Brake line temp sensor "C", L MLG BRK HTR LN 2 SYS 1&3 (V58T1702A)
final temp at 104 f deg.

Brake line temp sensor "D", LMLG BRK HTR LN 4 SYS 2&3 (V58T1703A)
final temp at 100 f deg.

Left outboard elevon actuator body (V58T0394) final temp at 108 f deg.
Left inboard elevon actuator body () final temp at 141 f deg.


Hydraulic resvoirs # 1 @ 178 f deg, #2 @ 169 f deg, #3 @ 141 f deg

Back to the subject at hand AGAIN a accident has damaged the foam

Any chance this will be a OUT if the next launch loses foam?

sorry we must not have repaired the damage properly although we did try?

"Bob Haller" wrote in message
oups.com...
Back to the subject at hand AGAIN a accident has damaged the foam

Any chance this will be a OUT if the next launch loses foam?

sorry we must not have repaired the damage properly although we did try?

I know I would be very concerned, but then, I've come to expect the worst
from our manned space program these days.

George

Youre concerns are shared by all, for the biggest concern isnt the
money it should be taking care of the astronauts. The subject at hand
is a safe shuttle flight, with all components of the orbiter system
working properly, external tank, and hydraulic system included.
Determining the fate of the orbiter, and our space program involves
more than just if foam will popcorn from the external tank, it is a
safe return of our astronauts.

With all these problems with gravity they seem to have over there, maybe
they need to tie helium balloons to everything!

:-)

Brian

--
Brian Gaff....Note, this account does not accept Bcc: email.
graphics are great, but the blind can't hear them
Email:
__________________________________________________ __________________________________________________ __________


"George" wrote in message
news:lZTWf.897982$xm3.574772@attbi_s21...
Not again!

http://www.cnn.com/2006/TECH/space/0...eut/index.html

Thursday, March 30, 2006; Posted: 10:19 a.m. EST (15:19 GMT)

CAPE CANAVERAL, Florida (Reuters) -- NASA is investigating another mishap
at the Kennedy Space Center, this time an accident involving the remodeled
fuel tank to be used for the next shuttle mission, the agency said on
Wednesday.
Technicians were replacing a vent valve near the top of the 154-foot
(47-meter) tall tank on Tuesday when a Halogen work lamp fell and hit the
tank's foam insulation.

Preliminary inspections show the impact left five small indentations, with
the largest about the size of a stick of gum, and one 6-inch
(15-centimeter) to 7-inch (17-centimeter) long scratch, said Marion
LaNasa, spokesman for tank manufacturer Lockheed Martin Corp.

A detailed inspection of the area was under way, LaNasa said, but the
incident was not expected to affect the shuttle's targeted July 1 liftoff.

The affected area is not among the sections of tank foam insulation that
were redesigned after the 2003 Columbia disaster and again after the July
2005 flight of Discovery, the only launch since the accident.

A piece of foam insulation that fell off the tank and hit Columbia's wing
during liftoff was responsible for heat shield damage that led to the
ship's destruction and the loss of seven crewmembers during atmospheric
re-entry on February 1, 2003.

A similar problem occurred during Discovery's liftoff 2 1/2 years later,
though the shuttle escaped damage.

NASA is preparing Discovery for launch again, but it first must prove that
the new tank design is safe to fly. A series of wind tunnel tests and
analyses are under way.

Safety has been a top priority for NASA, particularly at the shuttle
processing center in Florida, where a series of mishaps have resulted in a
death, equipment damage and several near-disasters over the past month.

I have to say...

That nothing that occurs today in processing will surprise me Hey
ooops they dropped a orbiter, ran a vehicle into one, had a VAB fire
with the solids could of taken out the building.

oh wait all this unbelievable stuff has already occured

George is right expect the worst so were nott disappointed...

"George" wrote in message
news:1l_Wf.68512$oL.5648@attbi_s71...
I know I would be very concerned, but then, I've come to expect the worst
from our manned space program these days.

Then you and Bob will get along famously in killfile hell.

Jeff
--
Remove icky phrase from email address to get a valid address.