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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 |
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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 @ |
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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 |
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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. |
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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 |
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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! [ ]=====================================[ |
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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 |
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NASA Astronaut on Columbia Repair (and others)
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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! [ ]=====================================[ |
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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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