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NASA Astronaut on Columbia Repair (and others)



 
 
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  #121  
Old November 19th 06, 03:22 PM posted to sci.space.history
columbiaaccidentinvestigation
external usenet poster
 
Posts: 1,344
Default NASA Astronaut on Columbia Repair (and others)

Pat Flannery wrote:" " well no i quess it isnt worth it, it was just
trolling fluff stuff not worth showing again.

And Yes pat, just for you next time ill try to include pictures for
your limited understanding, but until then troll somewhere else, it
just shows youre limited critical thinking, logic, and comprehension
skills... Oh yeah pat, besides calling me names (ie trolling) have you
provided anything to this thread, or do you think you usenet bullying
is productive.

Now back to the facts, as the rogers commission stated unequivocally
"Under the terms of the FRR Policy Directive, such damage would
appear to require discussion: "the scope of the review should cover
status and issues in areas such as . . . prior flight anomalies"
(*1), when referring to the previous flights (sts-61c) srbs evaluation
(*2) which clearly showed a burn through problem, but because of
attempting to meet schedule demands the time was not taken by mission
managers to make logical correlation of with the effects of cold
weather on the srb joint o-ring seal resiliency, when making a go for
launch decision in the cold weather the morning of jan 28, 1986.

Now implementing the recommendations contained in the diaz report to
the caib, in conjunction with a full quantitative risk assessment of
the shuttle system would provide nasa managers with the communication
structure, information, and technology to manage and understand the
technical input from others up and down the decision making process in
operating the shuttle safely within it's capabilities throughout the
fleets retirement process.

(*1)
http://www.gpoaccess.gov/challenger/64_420b.pdf
page (26 pdf) or 208 in the report
Post-flight disassembly of STS 61-C SRB hardware following its launch
on January 12 revealed that erosion of the primary O-ring had occurred
in the aft field joint of the left motor. Hot gas had also bypassed the
primary seal in the left nozzle joint. Erosion of the primary seal had
also occurred in the nozzle joint of the right motor.6 Under the terms
of the FRR Policy Directive, such damage would appear to require
discussion: "the scope of the review should cover status and issues
in areas such as . . . prior flight anomalies. . . ."
lies. . . ."

(*2)
And the sts-61c srb evaluation stated:
http://ntrs.nasa.gov/archive/nasa/ca...992075284..pdf
NSTS-22301, page 4
"SOLID ROCKET BOOSTER
The STS 61-C flight utilized lightweight solid rocket motor (SRM)
cases. SRM
propulsion performance was normal and within specification limits, with
propellant burn rates for both SRM's near predicted values. Solid
rocket booster (SRB) thrust differentials were within specification
throughout the flight....

A postflight evaluation of the SRM structure to determine the extent of
damage
revealed the following significant items:
a. A gas path was noted at the 154-degree position of the aft field
joint of the left S_M. Soot was found from the 140-degree to the
178-degree position, and soot was found in the primary groove from the
68-degree to the 183-degree (115 degrees arc) position. C-ring
damagewas noted at the 154-degree position with a maximumerosion depth
of 0.00_ inch and erosion length of 3.5 inches. The 0-ring was affected
by heat over a 14-inch length in this area.
b. A gas path was found from the 273.6-degree to the 309.6-degree (36
degrees arc) position of the left S_Mnozzle joint. Soot was found in
the primary 0-ring groove over the entire 360-degree circumference. A
potential impingement point was located at the 302.4-degree point;
however, no 0-ring damage was found.
c. A gas path was found at the 162-degree point with soot in the
primary 0-ring groove from the lOS-degree to the 220-degree (112
degrees arc) point on the right SRM nozzle joint. 0-ring damage was
found at the 162-degree point with the maximum erosion depth being
0.011 inch and the erosion length being 8 inches. The 0-rlng was
affected by heat over a 26-1nch length in this area.
d. A gas path was found on the outer surface of the igniter at the
130-degree point of the left SRM. Soot was found on the aft side of the
outer Gaskoseal, approaching the primary sea! over a 70-degree arc (130
to 200 degrees), and on the outer edge of the inner Gasko seal over a
130-degree arc (ii0 to 240 degrees), however, no seal damage was found.
e. A gas path was found on the outer surface of the igniter at the
250-degree point of the right S_. Soot was found on the inside edge of
the outer Gasko seal over the entire 360-degree circumference, however,
it did not progress beyond the edge of the seal. There was a slight
discoloration of the metal on both sides of the seal over the entire
360-degree circumference."

Probabilistic Risk Analysis for the NASA Space Shuttle:
A Brief History and Current Work by Elisabeth Paté-Cornell, Stanford
University,
and Robin Dillon, Virginia Tech Submitted for publication in
Reliability Engineering and System Safety April, 2000

Schedule pressures
Caib report vol 1 page 131, col 1, par 10
"Chapter 6, Decision Making at Nasa
Recomedations
6.2 SCHEDULE PRESSURE
Countdown to Space Station "Core Complete:" A Workforce Under
Pressure
During the course of this investigation, the Board received several
unsolicited comments from NASA personnel regard-ing pressure to meet a
schedule. These comments all con-cerned a date, more than a year after
the launch of Columbia, that seemed etched in stone: February 19, 2004,
the sched-uled launch date of STS-120. This flight was a milestone in
the minds of NASA management since it would carry a sec-tion of the
International Space Station called "Node 2." This would configure
the International Space Station to its "U.S. Core Complete" status.


Independent Technical Authority
Caib report vol 1 page 227 col 1
"Chapter 11 Recommendations
Organization
R7.5-1 Establish an independent Technical Engineering Authority that is
responsible for technical requirements and all waivers to them, and
will build a disciplined, systematic approach to identifying,
analyzing, and controlling hazards throughout the life cycle of the
Shuttle System. The independent technical authority does the fol-lowing
as a minimum:
· Develop and maintain technical standards for all Space Shuttle
Program projects and elements
· Be the sole waiver-granting authority for all technical standards
· Conduct trend and risk analysis at the sub-system, system, and
enterprise levels
· Own the failure mode, effects analysis and hazard reporting systems

· Conduct integrated hazard analysis
· Decide what is and is not an anomalous event
· Independently verify launch readiness
· Approve the provisions of the recertifica-tion program called for
in Recommendation R9.1-1.
The Technical Engineering Authority should be funded directly from NASA
Headquarters, and should have no connection to or responsibility for
schedule or program cost.
R7.5-2 NASA Headquarters Office of Safety and Mis-sion Assurance should
have direct line authority over the entire Space Shuttle Program safety
organization and should be independently re-sourced.
R7.5-3 Reorganize the Space Shuttle Integration Office to make it
capable of integrating all elements of the Space Shuttle Program,
including the Or-biter."


Open sharing of information is crucial to improving everybody's
understanding of the universe around us.
Tom

  #122  
Old November 19th 06, 04:16 PM posted to sci.space.shuttle,sci.space.history
Craig Fink
external usenet poster
 
Posts: 1,858
Default NASA Astronaut on Columbia Repair (and others)

On Sun, 19 Nov 2006 00:15:00 +0200, SENECA wrote:

I thought about. Maybe the surface flow at the RCC underside is
unimportant for the BLT. In front of the RCC is the bow shock,
what deflects some of the incomming flow. The wing is inclined
by 40 degree. Therefore the wing below the RCC gets part of the
flow from the free space in front of the shuttle. In other words,
at the middle of the wing near the wheel bay only a fraction of
the BL mass flow there had passed the RCC underside. I`m not sure
on that but it may one explanation.


I posted this on the flow on the bottom side of the Orbiter some time ago.
http://groups.google.com/group/sci.s...054dc0607262a4
Basically, the airflow in the boundary layer, turbulent or laminar, has
passed through a bow shock, and then is expanded as it flows around the
wing. Compressed in the shock wave to 10000F, much of the energy is then
used to decompose the air into O and N and the temperature decreases to
3000F before coming into contact with the RCC. As the air flows around the
wing, it's actually going through an expansion. Decreasing the temperature
of the O2, N2, O, N, NO plasma. This is the environment the tiles find
themselves in, across the entire bottom of the Orbiter.

To me there must be a hysteresis effect to the chemical reaction. The
actual time in the bow shock and the extreme high temperatures cause the
decomposition of the air to occur rather quickly. But, as the air flows
around the wing and is expanded, the opposite reaction where the O and N
recombines into O2, N2 and NOx is much much slower. So, the temperature of
the air in the boundary layer, turbulent or not, is within the temperature
range that the tiles can handle.

Near the middle of the wing and at the wheel bay, all the air in the
boundary layer has passed through a bow shock, decomposed, then as
been expanded significantly. For the nose of the Orbiter and directly
behind it, the expansion was 50 degrees. At the nose, the air in the
boundary layer has gone from zero velocity, thru Mach 1, and continued
acceleration up to hypersonic velocity as it is expanded 50 degrees around
the curved surface of the RCC nose cap.

So, with Columbia, part of the bow shock entered the wing. But the flow
around the hole that didn't, still had some expansion to go through. I
imagine that might have kept the boundary layer somewhat well behaved.

--
Craig Fink
Courtesy E-Mail Welcome @
  #123  
Old November 19th 06, 04:27 PM posted to sci.space.shuttle,sci.space.history
columbiaaccidentinvestigation
external usenet poster
 
Posts: 1,344
Default NASA Astronaut on Columbia Repair (and others)

Craig Fink wrote:"So, with Columbia, part of the bow shock entered
the wing. But the flow around the hole that didn't, still had some
expansion to go through. I imagine that might have kept the boundary
layer somewhat well behaved"


I find volume V part 13 very of the caibs report helpful, you might to
please see below.

REENTRY ANALYSIS:
Caib report vol v part 13 page 106 par 6
Chapter 4 Aerodynamics
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 189 par 2
Chapter 5 Aerothermodynamics
5.2.3.3 Results for Mach 6 Air
"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."


Caib report vol v part 13 page 305 par 2
Chapter 5 Aerothermodynamics
"5.3.2.5 Properties of Flow into Breach Holes
The engineering method was exercised for four hole sizes in leading
edge Panel 6 - 1, 2, 4, and 6 inches in diameter. Figure 5.3.2-5
presents the predicted fraction of the boundary layer that is ingested
into the hole Both the 4 and 6-inch diameter holes result in the
ingestion of the entire boundary layer at later times in the
trajectory. Figure 5.3.2-6 and Figure 5.3.2-7 present the predicted
bulk gas temperature and enthalpy of the ingested gas. The gas
temperatures are mostly between 9,000 and 10,000 R. Figure 5.3.2-8 and
Figure 5.3.2-9 present the predicted mass flow rate and energy flux and
compare the results to the values obtained from the CFD solutions for
the holes in Panel 6. For the smaller holes - 2 and 4-inch diameter
- the engineering method provides a very good estimate of the
ingested flow rate and energy flux. However, as the hole is increased
in size to the 6-inch diameter, the engineering methods tend to under
predict the values
Breach holes in Panel 8 were also analyzed with the engineering method.
In this case, the external pressure used for the computation (Cp=1.46)
was for a location slightly different from the location where the
curve-fits were obtained. Figure 5.3.2-10 and Figure 5.3.2-11 present
the predicted mass flux and energy flux into the holes. The values for
the CFD solution with a 10-inch diameter hole are also plotted and
indicate that the engineering method is under predicting the energy
ingested into large breach holes".


Caib report vol v part 13 page 311 par 3
Chapter 5 Aerothermodynamics
5.3.3.1 Basic internal plume physics and impingement issues
The nature and structure of a free jet issuing into the Orbiter
interior through a breach in the thermal protection system is dictated
by the total pressure difference across the surface, the hole diameter,
boundary layer properties, and local boundary layer edge Mach number.
For small orifices, on the order of one inch, the jet may be assumed to
enter normal to the internal surface with a sonic condition. For such a
case the governing
parameter that dictates shock structure and mixing of the jet is the
ratio of external riving pressure to internal pressure. Bulk fluid
properties are a function of the percentage of external boundary layer
drawn off into the hole. For very small hole, the properties will be
near wall conditions, whereas larger holes can produce internal flows
with enthalpy levels approaching free stream total values. As the
high-speed gas enters the cavity, it immediately starts mixing with the
ambient fluid at the boundaries of the jet. This mixing zone gets
larger as the jet progresses, until finally the core portion of the jet
has been consumed and the jet has reached a fully developed condition.
Depending on the conditions driving the jet, this may not occur until
ten's of diameters downstream. In the highly under-expanded state,
the jet shock structure is dominated by a normal shock downstream of
the initial expansion called the Mach disk. Immediately downstream of
the Mach disk the flow is subsonic, though it may re-expand to
supersonic flow. Figure 5.3.3-1 displays variations in free-jet
structure and with varying pressure ratios. Figure 5.3.3-2 shows the
impact of pressure ratios in the range expected for Orbiter
penetrations on computed flow structure. Larger hole diameters display
significant departure from this relatively simple structure as larger
percentages of the highly energetic boundary layer are ingested and
increased transverse momentum bends the jet over in the direction of
the boundary layer edge flow. This effect was discovered with the first
fully coupled,
internal/exterior flow solution performed for a two-inch breach into
RCC panel 6. Complete results are presented in Section 5.3.6.1.4.
Larger diameter penetrations tend to carry highly supersonic, high
temperature gases directly to the interior surfaces and produce highly
complex shock/impingement structures that can significantly impact
local heat transfer rates."


Caib report vol v part 13 page 535 par 7
Chapter 6 Thermal
6.7 Damaged Wing Leading Edge Coupled Aero-Thermal-Structural Analysis
"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) from the OEX flight data with the model
predictions at an analogous location on panel 8 (in this case panel 8
in the model is used as a surrogate for panel 9 as noted previously).
The average predicted temperature of two nodes on the rear facesheet
are used in the comparison for each damage case. Up to the flight
estimated time of spar breach at approximately 490 seconds the
predicted thermal response for the 6 inch case is in reasonable
agreement. After this point, the predicted temperature rise rates are
much slower than flight data, indicating the effect of convective
heating experienced during flight in this area from the hot gas jet
expanding into the wing interior. Modeling of such heating was not
included in this analysis."

Caib report vol V page 539, par 7
Chapter 6 Thermal
AeroAerothermalThermalStructuresTeamFinalReport
"6.11 Leading Edge Reinforced Carbon-Carbon (RCC) Hole Growth Thermal
Analysis
A prediction of RCC hole growth was performed using JSC arc jet test
data obtained from hypervelocity impacted RCC test specimens when
subject to a high temperature entry environment. The objective of the
arc jet testing was to establish the oxidation characteristics of RCC
with thru holes obtained from hypervelocity impacts. The specimens were
exposed to constant heating conditions at temperatures of 2500 and
2800F and pressures of 50 to 180 psf. Correlations were developed from
the data for use in trajectory simulations to predict hole growth and
hot gas flow through an enlarging hole into the wing leading edge
cavity.
A 0.75 inch diameter hole in the RCC was assumed for analysis purposes.
Figure 6-97 shows the heat flux and pressure environment at the hole
while Figure 6-98 shows the resulting RCC surface temperature as a
function of time. The predicted RCC temperature of approximately
4800°F is assumed to be consistent with a diffusion-limited erosion
regime for bare or uncoated RCC. With this assumption, the erosion or
hole growth rate measured for the 2800°F arc jet tests can be used for
erosion rate estimates here. The erosion rate in this flight
environment and regime is .0032 in/sec. Figure 6-99 reveals the results
of the analysis and shows the predicted growth to a final OML diameter
of 4.0 inches. The predicted IML (back-face) diameter is slightly
smaller at 3.0 inches. Extrapolation of this analysis to higher RCC
temperatures (sublimation regime) or larger initial hole diameter is
not recommended since the data base is very limited."


Open sharing of information is crucial to improving everybody's
understanding of the universe around us.
Tom

  #124  
Old November 19th 06, 04:45 PM posted to sci.space.history
Scott Hedrick
external usenet poster
 
Posts: 724
Default NASA Astronaut on Columbia Repair (and others)


"OM" wrote in message
...
On Sat, 18 Nov 2006 08:50:55 -0500, "Scott Hedrick"
wrote:

"columbiaaccidentinvestigation"
wrote in message
groups.com...

Open sharing of information is crucial to improving everybody's
understanding of the universe around us.


Even more crucial is the *acceptance of information that conflicts with
your
beliefs*. Hint hint.


...If it smells like a Znkfba, and trolls like a Znkfba, it probably
*is* a Znkfba.


Which is why, after reading his reply to my statement above, I killfiled
him.


  #125  
Old November 19th 06, 04:53 PM posted to sci.space.history
columbiaaccidentinvestigation
external usenet poster
 
Posts: 1,344
Default NASA Astronaut on Columbia Repair (and others)

scott said in clinging to his ignorance: "Which is why, after reading
his reply to my statement above, I killfiled him."

Just read (scott you can do it), as I find volume V part 13 very
helpful you might as well so please see below.

REENTRY ANALYSIS:
Caib report vol v part 13 page 106 par 6
Chapter 4 Aerodynamics
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 189 par 2
Chapter 5 Aerothermodynamics
5.2.3.3 Results for Mach 6 Air
"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."


Caib report vol v part 13 page 305 par 2
Chapter 5 Aerothermodynamics
"5.3.2.5 Properties of Flow into Breach Holes
The engineering method was exercised for four hole sizes in leading
edge Panel 6 - 1, 2, 4, and 6 inches in diameter. Figure 5.3.2-5
presents the predicted fraction of the boundary layer that is ingested
into the hole Both the 4 and 6-inch diameter holes result in the
ingestion of the entire boundary layer at later times in the
trajectory. Figure 5.3.2-6 and Figure 5.3.2-7 present the predicted
bulk gas temperature and enthalpy of the ingested gas. The gas
temperatures are mostly between 9,000 and 10,000 R. Figure 5.3.2-8 and
Figure 5.3.2-9 present the predicted mass flow rate and energy flux and
compare the results to the values obtained from the CFD solutions for
the holes in Panel 6. For the smaller holes - 2 and 4-inch diameter
- the engineering method provides a very good estimate of the
ingested flow rate and energy flux. However, as the hole is increased
in size to the 6-inch diameter, the engineering methods tend to under
predict the values
Breach holes in Panel 8 were also analyzed with the engineering method.
In this case, the external pressure used for the computation (Cp=1.46)
was for a location slightly different from the location where the
curve-fits were obtained. Figure 5.3.2-10 and Figure 5.3.2-11 present
the predicted mass flux and energy flux into the holes. The values for
the CFD solution with a 10-inch diameter hole are also plotted and
indicate that the engineering method is under predicting the energy
ingested into large breach holes".


Caib report vol v part 13 page 311 par 3
Chapter 5 Aerothermodynamics
5.3.3.1 Basic internal plume physics and impingement issues
The nature and structure of a free jet issuing into the Orbiter
interior through a breach in the thermal protection system is dictated
by the total pressure difference across the surface, the hole diameter,
boundary layer properties, and local boundary layer edge Mach number.
For small orifices, on the order of one inch, the jet may be assumed to
enter normal to the internal surface with a sonic condition. For such a
case the governing
parameter that dictates shock structure and mixing of the jet is the
ratio of external riving pressure to internal pressure. Bulk fluid
properties are a function of the percentage of external boundary layer
drawn off into the hole. For very small hole, the properties will be
near wall conditions, whereas larger holes can produce internal flows
with enthalpy levels approaching free stream total values. As the
high-speed gas enters the cavity, it immediately starts mixing with the
ambient fluid at the boundaries of the jet. This mixing zone gets
larger as the jet progresses, until finally the core portion of the jet
has been consumed and the jet has reached a fully developed condition.
Depending on the conditions driving the jet, this may not occur until
ten's of diameters downstream. In the highly under-expanded state,
the jet shock structure is dominated by a normal shock downstream of
the initial expansion called the Mach disk. Immediately downstream of
the Mach disk the flow is subsonic, though it may re-expand to
supersonic flow. Figure 5.3.3-1 displays variations in free-jet
structure and with varying pressure ratios. Figure 5.3.3-2 shows the
impact of pressure ratios in the range expected for Orbiter
penetrations on computed flow structure. Larger hole diameters display
significant departure from this relatively simple structure as larger
percentages of the highly energetic boundary layer are ingested and
increased transverse momentum bends the jet over in the direction of
the boundary layer edge flow. This effect was discovered with the first
fully coupled,
internal/exterior flow solution performed for a two-inch breach into
RCC panel 6. Complete results are presented in Section 5.3.6.1.4.
Larger diameter penetrations tend to carry highly supersonic, high
temperature gases directly to the interior surfaces and produce highly
complex shock/impingement structures that can significantly impact
local heat transfer rates."


Caib report vol v part 13 page 535 par 7
Chapter 6 Thermal
6.7 Damaged Wing Leading Edge Coupled Aero-Thermal-Structural Analysis
"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) from the OEX flight data with the model
predictions at an analogous location on panel 8 (in this case panel 8
in the model is used as a surrogate for panel 9 as noted previously).
The average predicted temperature of two nodes on the rear facesheet
are used in the comparison for each damage case. Up to the flight
estimated time of spar breach at approximately 490 seconds the
predicted thermal response for the 6 inch case is in reasonable
agreement. After this point, the predicted temperature rise rates are
much slower than flight data, indicating the effect of convective
heating experienced during flight in this area from the hot gas jet
expanding into the wing interior. Modeling of such heating was not
included in this analysis."

Caib report vol V page 539, par 7
Chapter 6 Thermal
AeroAerothermalThermalStructuresTeamFinalReport
"6.11 Leading Edge Reinforced Carbon-Carbon (RCC) Hole Growth Thermal
Analysis
A prediction of RCC hole growth was performed using JSC arc jet test
data obtained from hypervelocity impacted RCC test specimens when
subject to a high temperature entry environment. The objective of the
arc jet testing was to establish the oxidation characteristics of RCC
with thru holes obtained from hypervelocity impacts. The specimens were
exposed to constant heating conditions at temperatures of 2500 and
2800F and pressures of 50 to 180 psf. Correlations were developed from
the data for use in trajectory simulations to predict hole growth and
hot gas flow through an enlarging hole into the wing leading edge
cavity.
A 0.75 inch diameter hole in the RCC was assumed for analysis purposes.
Figure 6-97 shows the heat flux and pressure environment at the hole
while Figure 6-98 shows the resulting RCC surface temperature as a
function of time. The predicted RCC temperature of approximately
4800°F is assumed to be consistent with a diffusion-limited erosion
regime for bare or uncoated RCC. With this assumption, the erosion or
hole growth rate measured for the 2800°F arc jet tests can be used for
erosion rate estimates here. The erosion rate in this flight
environment and regime is .0032 in/sec. Figure 6-99 reveals the results
of the analysis and shows the predicted growth to a final OML diameter
of 4.0 inches. The predicted IML (back-face) diameter is slightly
smaller at 3.0 inches. Extrapolation of this analysis to higher RCC
temperatures (sublimation regime) or larger initial hole diameter is
not recommended since the data base is very limited."


Open sharing of information is crucial to improving everybody's
understanding of the universe around us.
Tom

  #126  
Old November 19th 06, 06:55 PM posted to sci.space.history
OM[_4_]
external usenet poster
 
Posts: 806
Default NASA Astronaut on Columbia Repair (and others)

On Sun, 19 Nov 2006 10:45:10 -0500, "Scott Hedrick"
wrote:

...If it smells like a Znkfba, and trolls like a Znkfba, it probably
*is* a Znkfba.


Which is why, after reading his reply to my statement above, I killfiled
him.


....Same here. IIRC, this *might* be the same troll that harassed me a
couple of years ago about the CLFAQ, demanding that I include a
section about conspiracy and/or ineptitude amongst the various
managers and contractors. The whole thing smacked of Znkfba's rants
against Morton Thiokol, so I told him where to go and blocked e-mail
from his account.


OM
--
]=====================================[
] OMBlog - http://www.io.com/~o_m/omworld [
] Let's face it: Sometimes you *need* [
] an obnoxious opinion in your day! [
]=====================================[
  #127  
Old November 19th 06, 07:02 PM posted to sci.space.history
columbiaaccidentinvestigation
external usenet poster
 
Posts: 1,344
Default NASA Astronaut on Columbia Repair (and others)

OM wrote:
On Sun, 19 Nov 2006 10:45:10 -0500, "Scott Hedrick"
wrote:

...If it smells like a Znkfba, and trolls like a Znkfba, it probably
*is* a Znkfba.


Which is why, after reading his reply to my statement above, I killfiled
him.


...Same here. IIRC, this *might* be the same troll that harassed me a
couple of years ago about the CLFAQ, demanding that I include a
section about conspiracy and/or ineptitude amongst the various
managers and contractors. The whole thing smacked of Znkfba's rants
against Morton Thiokol, so I told him where to go and blocked e-mail
from his account.



Yes om, you too can cling to youre ignorance as well it is your own
free will, or just read the information, as you have no way of refuting
me other than personal attacks it is obvious youre actions are that of
a troll, trolling for bites.

Once again, I find volume V part 13 very helpful you might to please
see below.

REENTRY ANALYSIS:
Caib report vol v part 13 page 106 par 6
Chapter 4 Aerodynamics
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 189 par 2
Chapter 5 Aerothermodynamics
5.2.3.3 Results for Mach 6 Air
"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."


Caib report vol v part 13 page 305 par 2
Chapter 5 Aerothermodynamics
"5.3.2.5 Properties of Flow into Breach Holes
The engineering method was exercised for four hole sizes in leading
edge Panel 6 - 1, 2, 4, and 6 inches in diameter. Figure 5.3.2-5
presents the predicted fraction of the boundary layer that is ingested
into the hole Both the 4 and 6-inch diameter holes result in the
ingestion of the entire boundary layer at later times in the
trajectory. Figure 5.3.2-6 and Figure 5.3.2-7 present the predicted
bulk gas temperature and enthalpy of the ingested gas. The gas
temperatures are mostly between 9,000 and 10,000 R. Figure 5.3.2-8 and
Figure 5.3.2-9 present the predicted mass flow rate and energy flux and
compare the results to the values obtained from the CFD solutions for
the holes in Panel 6. For the smaller holes - 2 and 4-inch diameter
- the engineering method provides a very good estimate of the
ingested flow rate and energy flux. However, as the hole is increased
in size to the 6-inch diameter, the engineering methods tend to under
predict the values
Breach holes in Panel 8 were also analyzed with the engineering method.
In this case, the external pressure used for the computation (Cp=1.46)
was for a location slightly different from the location where the
curve-fits were obtained. Figure 5.3.2-10 and Figure 5.3.2-11 present
the predicted mass flux and energy flux into the holes. The values for
the CFD solution with a 10-inch diameter hole are also plotted and
indicate that the engineering method is under predicting the energy
ingested into large breach holes".


Caib report vol v part 13 page 311 par 3
Chapter 5 Aerothermodynamics
5.3.3.1 Basic internal plume physics and impingement issues
The nature and structure of a free jet issuing into the Orbiter
interior through a breach in the thermal protection system is dictated
by the total pressure difference across the surface, the hole diameter,
boundary layer properties, and local boundary layer edge Mach number.
For small orifices, on the order of one inch, the jet may be assumed to
enter normal to the internal surface with a sonic condition. For such a
case the governing
parameter that dictates shock structure and mixing of the jet is the
ratio of external riving pressure to internal pressure. Bulk fluid
properties are a function of the percentage of external boundary layer
drawn off into the hole. For very small hole, the properties will be
near wall conditions, whereas larger holes can produce internal flows
with enthalpy levels approaching free stream total values. As the
high-speed gas enters the cavity, it immediately starts mixing with the
ambient fluid at the boundaries of the jet. This mixing zone gets
larger as the jet progresses, until finally the core portion of the jet
has been consumed and the jet has reached a fully developed condition.
Depending on the conditions driving the jet, this may not occur until
ten's of diameters downstream. In the highly under-expanded state,
the jet shock structure is dominated by a normal shock downstream of
the initial expansion called the Mach disk. Immediately downstream of
the Mach disk the flow is subsonic, though it may re-expand to
supersonic flow. Figure 5.3.3-1 displays variations in free-jet
structure and with varying pressure ratios. Figure 5.3.3-2 shows the
impact of pressure ratios in the range expected for Orbiter
penetrations on computed flow structure. Larger hole diameters display
significant departure from this relatively simple structure as larger
percentages of the highly energetic boundary layer are ingested and
increased transverse momentum bends the jet over in the direction of
the boundary layer edge flow. This effect was discovered with the first
fully coupled,
internal/exterior flow solution performed for a two-inch breach into
RCC panel 6. Complete results are presented in Section 5.3.6.1.4.
Larger diameter penetrations tend to carry highly supersonic, high
temperature gases directly to the interior surfaces and produce highly
complex shock/impingement structures that can significantly impact
local heat transfer rates."


Caib report vol v part 13 page 535 par 7
Chapter 6 Thermal
6.7 Damaged Wing Leading Edge Coupled Aero-Thermal-Structural Analysis
"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) from the OEX flight data with the model
predictions at an analogous location on panel 8 (in this case panel 8
in the model is used as a surrogate for panel 9 as noted previously).
The average predicted temperature of two nodes on the rear facesheet
are used in the comparison for each damage case. Up to the flight
estimated time of spar breach at approximately 490 seconds the
predicted thermal response for the 6 inch case is in reasonable
agreement. After this point, the predicted temperature rise rates are
much slower than flight data, indicating the effect of convective
heating experienced during flight in this area from the hot gas jet
expanding into the wing interior. Modeling of such heating was not
included in this analysis."

Caib report vol V page 539, par 7
Chapter 6 Thermal
AeroAerothermalThermalStructuresTeamFinalReport
"6.11 Leading Edge Reinforced Carbon-Carbon (RCC) Hole Growth Thermal
Analysis
A prediction of RCC hole growth was performed using JSC arc jet test
data obtained from hypervelocity impacted RCC test specimens when
subject to a high temperature entry environment. The objective of the
arc jet testing was to establish the oxidation characteristics of RCC
with thru holes obtained from hypervelocity impacts. The specimens were
exposed to constant heating conditions at temperatures of 2500 and
2800F and pressures of 50 to 180 psf. Correlations were developed from
the data for use in trajectory simulations to predict hole growth and
hot gas flow through an enlarging hole into the wing leading edge
cavity.
A 0.75 inch diameter hole in the RCC was assumed for analysis purposes.
Figure 6-97 shows the heat flux and pressure environment at the hole
while Figure 6-98 shows the resulting RCC surface temperature as a
function of time. The predicted RCC temperature of approximately
4800°F is assumed to be consistent with a diffusion-limited erosion
regime for bare or uncoated RCC. With this assumption, the erosion or
hole growth rate measured for the 2800°F arc jet tests can be used for
erosion rate estimates here. The erosion rate in this flight
environment and regime is .0032 in/sec. Figure 6-99 reveals the results
of the analysis and shows the predicted growth to a final OML diameter
of 4.0 inches. The predicted IML (back-face) diameter is slightly
smaller at 3.0 inches. Extrapolation of this analysis to higher RCC
temperatures (sublimation regime) or larger initial hole diameter is
not recommended since the data base is very limited."


Open sharing of information is crucial to improving everybody's
understanding of the universe around us.
Tom

  #128  
Old November 19th 06, 08:06 PM posted to sci.space.shuttle,sci.space.history
Jorge R. Frank
external usenet poster
 
Posts: 2,089
Default NASA Astronaut on Columbia Repair (and others)

wrote in
:

The risk of micrometeoroid or debris damage to the RCC panels has
been evaluated several times. Hypervelocity impact testing, using
nylon, glass, and aluminum projectiles, as well as low-velocity
impact testing with ice, aluminum, steel, and lead projectiles,
resulted in the addition of a 0.03- to 0.06-inch-thick layer of
Nextel-440 fabric between the Inconel foil and Cerachrome
insulation. Analysis of the design change predicts that the Orbiter
could survive re-entry with a quarter-inch diameter hole in the
lower surfaces of RCC panels 8 through 10 or with a one-inch hole in
the rest of the RCC panels.

This reffers to a time before CAIB. The last sentence is only possible
if they did extensive testing in the arcjet facility. Together with
the other statments by Madden it rules out any importance of the
oxidation at the crack surface in the Columbia case. The CAIB
conclusions still stands.


You are confusing the issue. I do not take issue with the CAIB's
conclusion regarding the actual Columbia working scenario. My issue is
with the hypothetical repair scenario.

The 2004 arcjet tests were higher fidelity than previous tests and
therefore invalidate the earlier test results you reference here. Prior
to the 2004 tests it was felt that the Nextel fabric would be adequate by
itself for the damage sizes quoted and that those sizes could be
established as the upper bound of the "No repair necessary" threshold.
After the 2004 tests, it was realized that this threshold needed to be
much lower.

You got your idea of the overlooked issue of RCC oxidation after you
heard something and presented your own conclusions from it here some
months ago.


Incorrect. I knew about the RCC oxidation issue since I saw the
presentation charts for the arcjet test results in 2004. That was long
before I posted about it on Usenet.

Ok, let us look in this example. Show me the source. I`m very
interested in the details of this test.


I'm sure you are. But as far as I know, the presentation is
unpublished, and I have no intention of jumping through the
export-control hoops it would take to publish it.


If this guys hide it behind arms trade regulations better you no
longer trust em in all things.


"This guys" are not hiding anything. The shuttle is on the ITAR Munitions
List and all shuttle technical publications must be pre-cleared to ensure
that sensitive technologies are not compromised. The presentation in
question does not necessarily contain such information, but it has not
received the necessary Export Control pre-clearance and I am not about to
do that for the sake of a silly argument on Usenet.

Maybe it was a presentation intended to
mind wash some people.


Highly doubtful. There were no references to Columbia at all in the
presentation and there is no indication that the presenters had any
motivation other than future RCC repair capability. The connection
between the results of the presentation and a hypothetical Columbia
repair scenario are entirely my own.

--
JRF

Reply-to address spam-proofed - to reply by E-mail,
check "Organization" (I am not assimilated) and
think one step ahead of IBM.
  #129  
Old November 19th 06, 08:13 PM posted to sci.space.history
OM[_4_]
external usenet poster
 
Posts: 806
Default NASA Astronaut on Columbia Repair (and others)

On Sun, 19 Nov 2006 13:06:47 -0600, "Jorge R. Frank"
wrote:

Incorrect. I knew about the RCC oxidation issue since I saw the
presentation charts for the arcjet test results in 2004. That was long
before I posted about it on Usenet.


....What? You were holding out on us, Jorge?? For shame! :-) :-) :-)

OM
--
]=====================================[
] OMBlog - http://www.io.com/~o_m/omworld [
] Let's face it: Sometimes you *need* [
] an obnoxious opinion in your day! [
]=====================================[
  #130  
Old November 19th 06, 09:01 PM posted to sci.space.shuttle,sci.space.history
columbiaaccidentinvestigation
external usenet poster
 
Posts: 1,344
Default NASA Astronaut on Columbia Repair (and others)

Jorge R. Frank wrote: "I knew about the RCC oxidation issue since I saw
the presentation charts for the arcjet test results in 2004. That was
long before I posted about it on Usenet"

Here is a link to a study of rcc oxidation conducted in 20000 titled
"NASA/TP-2000-209760 Oxidation of Reinforced Carbon-Carbon
Subjected to Hypervelocity Impact"

http://ston.jsc.nasa.gov/collections...000-209760.pdf

Open sharing of information is crucial to improving everybody's
understanding of the universe around us.
Tom

 




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