Ablative Coatings

During the early 1960s, major aerospace contractors during the early preconcept phases of space shuttle development were becoming increasingly interested in silicone-based elastomeric ablative coatings as possible heat shields. Engineers believed this type of ablator offered several advantages over the resin ablators used on previous capsules, including ease of application to complex shapes; flexibility over a wide range of temperatures; potential for refurbishment with spray, bonded sheets, or prefabricated panels; and superior shielding effectiveness at low-to – moderate heating rates. This coating would have to be a good insulator, lightweight, and easy to apply, remove, and reapply before another flight. The first real-world opportunity to test the materials on a full-scale reusable vehicle would come on X-15A-2 during its envelope expansion to Mach 8.[295

It was obvious that the Mach 6.5 structural design of the X-15 was not adequate to handle the aerodynamic heating loads expected at Mach 8. For example, the total heat load for a location on the underside of the nose was approximately 2,300 Btu per square foot at Mach 6, but over 13,000 Btu at Mach 8. Similarly, the wing leading edge absorbed 9,500 Btu per square foot at Mach 6, but 27,500 Btu at Mach 8. It might have been possible to beef up the hot structure to accommodate these heat loads, but this would have amounted to an extensive redesign the program could not afford.[296]

Researchers believed the ability of the ablator to protect the airplane might well be the governing factor during the envelope expansion. To provide an engineering tool to evaluate this problem during the planning of these flights, the AFFTC developed a real-time temperature simulation using the former Dyna-Soar hybrid simulator. In conjunction with a complete fixed-base simulation of X-15A-2, the hybrid had ability to predict the temperature at selected points for both protected and unprotected surfaces. Researchers obtained a temperature-time history from these simulations for a point aft of the nose-gear door for a flight to Mach 7.6 at 100,000 feet. They then compared this with the temperature at the same location for an actual Mach 6 flight. Both the effective heating rate and the maximum temperature were significantly more severe at the higher speed.-1297

There had been some minor interest in the use of ablators for the X-15 as early as 1961. For instance, on flight 1-23-39 researchers tested a sample of Avcoat no. 2 on the leading edge of the right wing, directly over the semispan thermocouple. The leading-edge temperature at 144 seconds after launch was only 25°F underneath the test sample, and the thermocouple on either side of it showed 350°F and 315°F. Nevertheless, since the entire point of the X-15 was to gather accurate aero-thermo data, it made no sense to protect the structure, until now.-298

It appears that the ablator initially chosen by North American for X-15A-2 was Emerson Electric Thermolag 500, and this is the product shown in most reference documentation as late as the end of 1964. North American extensively tested this material in its 2.5-inch, 1-megawatt plasma tunnel for up to 317 seconds at a time, even though only 180 seconds were required for the actual X-15A-2 flight conditions. The material thickness on the leading edge was 0.70 inch, the forward fuselage ranged between 0.20 inch and 0.04 inch, and the wing mid-span quarter-chord thickness was 0.10 inch. A commercial paint spray gun applied the material, which weighed only 303 pounds.-298

After further evaluation, however, researchers decided the material was unacceptable, primarily because of its cure cycle. The coating had to be subjected to 300°F for a prolonged period to cure properly, and although this had not been a serious problem for small test areas, accomplishing it on the entire airplane would have been a challenge. In addition, researchers found that T-500 was somewhat water-soluble after it cured-not an ideal trait for something that was to be used outdoors, even in the high desert.298

In late 1963 the Air Force and NASA formed a joint committee to select a more suitable ablative material, although T-500 continued as the baseline for another year. To determine which ablative materials qualified as candidates for use on the X-15, the committee set up an evaluation program and requested all major ablator manufacturers to provide test samples. The primary factors used in evaluating the materials were the shielding effectiveness, room-temperature cure cycle, bond integrity, operational compatibility with the X-15, and refurbishment. The researchers used three facilities for this evaluation, including the 2-inch arc jet tunnel at the University of Dayton Research Institute, the 2.5-megawatt arc tunnel at Langley, and the X-15 airplanes. They ranked the materials in order of their shielding effectiveness as measured under a low heat-flux environment, and sent the results to the Air Force Materials Laboratory at Wright-Patterson

afb.297

While North American was rebuilding the second airplane, NASA began initial flight tests of various ablative coatings on X-15-1 and X-15-3. Engineers applied the coatings to removable panels behind the ball nose, and directly to locations under the liquid-oxygen tank, on the lower surface of the horizontal stabilizers, and on the canopy, ventral stabilizer, speed brakes, and

rudder. The ventral stabilizer and speed brakes provided moderate heating rates in easily accessible locations that could tolerate material failures if they occurred. The liquid-oxygen tank provided a test area for checking the bond integrity at temperatures approaching -300°F during actual flight. The removable nose panels provided measured back-surface temperatures and allowed direct comparison of two materials under the same heating conditions. Researchers expected the canopy application to show whether a windshield-contamination problem existed, but the tests proved inconclusive.13021

Flight-testing began in late 1963 and concluded in October 1964. NASA wanted to find a material that could provide protection at heating rates of 5-150 Btu per square foot per second and shearing stresses as high as 15 psf at a total weight of less than 400 pounds. The bonding had to be reliable at skin temperatures from -300 to +500°F, and ideally the material should not require special curing or handling.13031

Eventually, 15 different materials were flight-tested and the more promising included General Electric ESM 1004B, Martin MA-32H and MA-45R, McDonnell B-44, and NASA E-2A-1 Purple Blend. Researchers at Langley were developing the NASA material primarily as a backup in case the commercial products did not prove acceptable. The evaluation group also performed limited tests of alternate forms of the Martin and McDonnell materials, and ultimately selected one of these, MA-25S, for full-scale use.13041

Flight-testing proved to be an extremely valuable part of the overall evaluation. Researchers discovered numerous deficiencies in materials, bond systems, and spray techniques during the flights that they probably would not have found any other way-another example of the fact that there is no substitute for real-world experience. The flight conditions experienced at Mach 5 showed material problems that had not appeared in ground-facility tests, mainly poor bonding and excessive erosion and blistering on some segments.13051

Most of these problems, if they had occurred during a Mach 8 flight, would have likely resulted in the loss of the airplane. One of the most serious problems was bond failures of sheet materials, usually because the material was too stiff to conform to skin irregularities, resulting in voids in the bond (glue). This proved to be a major blow to the concept of using ablators, since researchers had expected to be able to easily service the sheet materials before and after flight. The alternative was to apply the ablator with a spray gun, but many of the materials responded by delaminating and peeling off during flight. In every case examined in detail, this was the result of improper application, not a material failure. Nevertheless, it pointed out the difficulties of actually using these materials, and the test areas were generally only a couple of square feet-imagine the problems involved with coating an entire airplane.13061

CALCULATED TEMPERATURES

UNPROTECTED X-15-2 AIRFRAME

MAXIMUM VELOCITY = 3,000 fp*

ALTITUDE = 100r000 fl

2400

1602

1171

03

1082

1500

UNPROTECTED

1000

TYPICAL X-15

TEMPERATURE.

Г5700 IpiH .■

so о

ABLATIVE

CQATID

50 100 ISO 200 2S0 ЭОО 35 □ 400

TIME FROM LAUNCH, wt

The Mach 6.5 structural design of the original X-15 was not adequate to handle the aerodynamic heating loads expected at Mach 8 for the advanced X-15A-2. For example, the total heat load for a location on the underside of the nose was approximately 2,300 Btu per square foot at Mach 6, but over 13,000 Btu at Mach 8. Similarly, the wing leading edge absorbed 9,500 Btu per square foot at Mach 6, but 27,500 Btu at Mach 8. To protect the airframe, researchers turned to ablative coatings similar to ones being proposed for the space shuttle. (NASA)

A few materials eroded very badly on the ventral stabilizer leading edge. This was a sign of inadequate thermal protection since Mach 5 provided a low heating environment compared to the expected Mach 8 design requirements. For instance, the test panel under the nose reached a peak surface temperature of 1,000°F on a Mach 5 mission; at Mach 8, this panel would soar to 1,750°F.[307]

Something all the materials had in common was that they were difficult to remove after flight.

Char and remaining virgin material required soaking in solvents and manual scraping. One alternative that was tested was applying pressure-sensitive tape to the airframe, and then

applying the ablative over the tape. Technicians would simply strip the tape off after a flight and all residual material would come off with it, leaving a clean surface. However, if the tape got too hot-even in small areas-it could start to peel, taking the ablator with it, and leaving the airframe exposed to catastrophic heating levels.-1308!

As the flight-testing was nearing completion, researchers began thermal-performance testing using the 2.5-megawatt arc tunnel at Langley to determine the relative shielding effectiveness of the candidate materials. These tests closely simulated the peak heating rates and enthalpy levels expected on the design Mach 8 mission. The material manufacturers provided test samples of their materials installed on identical leading-edge and afterbody models.!3001

The leading-edge tests showed that most of the silicone-based ablators were unable to withstand the severe heating conditions. The three silicone-based materials had densities between 32 and 60 pounds per cubic foot, resulting in a surface between 0.545 inch and 0.294 inch thick. The back surface temperature of all three products was relatively similar, but the materials experienced a variety of erosion, blistering, and cracking problems during the tests. The fourth material tested in the Langley facility was a phenolic-silica ablator with a density of 110 pounds per cubic foot, resulting in a surface thickness of only 0.165 inch. The shape retention of this material was excellent, but its shielding effectiveness was low. All four of the materials passed the afterbody tests, with no significant differences in performance noted.13301

During the arc-tunnel tests, researchers observed that loosened material from the ablator tended to reattach to surfaces downstream. Flight tests on X-15-1 with a panel of windshield glass mounted on the vertical stabilizer aft of a sample patch of the ablator showed that the glass panel quickly became opaque, which would seriously restrict the pilot’s vision. Since the pilot obviously needed to see during landing, researchers considered three different approaches to restore the necessary vision. These included explosive fragmentation of the outer windshield glass after the high-speed run was completed, boundary-layer blowing over the windshield during the entire flight, and a hinged metal "eyelid" that could be opened after the high-speed portion of the flight.13111

The explosive concept worried everybody and was not pursued very far because there seemed to be too many possible failure modes. The boundary-layer idea was the only one that potentially provided a continuously clear windshield; however, the pilot actually had little reason to need completely clear vision at 100,000 feet since there was really nothing to run into at that altitude, and the implementation was complex and expensive. Therefore, the program selected the eyelid because it was the easiest to implement. The right windshield was unprotected and provided normal pilot vision during launch and initial climb-out. During the high-speed run, the right windshield would become opaque, allowing the pilot to see little more than light and dark patches of sky. The eyelid was installed over the left windshield; it would remain closed during the climb – out and high-speed flight, and open once the airplane slowed below Mach 3. The pilot would look out of the left side of the windshield for landing. This carried some risks, though. After one of his windshields shattered during a 1961 flight, Bob White reported that his vision had been "compromised" during landing. When flight tests began, the pilots discovered another phenomenon: the open eyelid created a small canard effect, causing the airplane to pitch up, roll right, and yaw right. The effects were small but noticeable.13121

In the end, the Air Force and NASA determined that the General Electric, Martin, McDonnell, and NASA Purple Blend products were all potentially acceptable and sent requests for proposals to the manufacturers. The source evaluation board received the proposals during late 1965, and in January 1966, NASA awarded a contract to Martin Marietta to design and apply a sprayable ablator

to X-15A-2.13131

The basic MA-25S ablative material had a virgin material density of 28 pounds per cubic foot. Martin had developed MA-25S "specifically for application over complex vehicle configurations," although it had existed well before the X-15 application was proposed. Most significantly, application and curing took place at room temperature (70°F to 100°F). A special premolded fiber-reinforced elastomeric silicone material (ESA-3560-IIA) similar to that used on the Air Force X-23A PRIME reentry vehicles would cover all the leading edges. Martin developed a premolded flexible material (MA-25S-1) to cover the seams around access panels, and used smaller pieces of this material to cover fasteners and other items that required last-minute access.13141

Interestingly, although Martin considered MA-25S a "mature" product, "all previous applications had been accomplished with laboratory equipment," and in March 1966 the company had to start from scratch to come up with methods to coat an entire airplane. Once the engineers finished writing the procedure, Martin procured several large sheets of Inconel and used them as test subjects. The company also ran compatibility tests with the various liquids and gases found on the X-15. Hydraulic fluid, helium, nitrogen, and ammonia did not seem to present any problems. An outside laboratory had to test the hydrogen peroxide, delaying the results, but no problems were expected. However, the MA-25S material, like all of the ablators originally tested for the X – 15, was impact-sensitive after exposure to liquid oxygen. Tests showed that a local detonation would occur on the material if it was submerged in liquid oxygen and struck with a force as low as 8 foot-pounds. Martin concluded that "the significance of the material being impact sensitive with liquid oxygen is not well understood at this time and this particular material characteristic should be reviewed with X-15A-2 operations personnel."3151

The sensitivity to liquid oxygen brought several unexpected problems since the casual spilling of liquid oxygen (not an uncommon occurrence) suddenly became a major problem. In response, Martin proposed spraying a white protective wear layer over the ablator to isolate it from any minor liquid-oxygen spillage. Nevertheless, the potential for contaminating the inside of the liquid-oxygen lines, pumps, vents, etc. during the application (spraying and sanding) of the ablator was the most worrisome.-13161

On 18 May 1966, X-15A-2 flight 2-44-79 provided the first relatively large-scale tests of MA – 25S and the ESA-3560-IIA leading-edge material. The materials had been applied (as appropriate) to three nose panels (F-3, F-4, and E-4), the UHF antenna, both main landing skids and struts, both sides of the ventral stabilizer, both lower speed brakes, and the left horizontal stabilizer. Researchers instrumented all of these panels to determine the effects of the ablator. Ground handling resulted in ablator damage that technicians repaired using a documented repair procedure; the test would inadvertently provide validation of its reparability. As part of the evaluation, technicians used various application techniques in different locations, providing some validation of the proposed concepts. In general, these tests were successful, although instrumentation failures precluded the gathering of any precise data from the nose panels.13171

The ablator also forced the program to develop a new pitot-static system. NASA relocated the static pickups since ablative material now covered the normal locations on the sides of the forward fuselage. Engineers moved the static source into a vented compartment behind the canopy that tests on X-15-1 had shown to be acceptable. An extendable pitot tube replaced the standard dogleg pitot ahead of the canopy because the temperatures expected at Mach 8 would exceed the standard tube’s limits. The retractable tube would remain within the fuselage until the aircraft decelerated below Mach 2, at which point the pilot would actuate a release mechanism and the tube would extend into the airstream. This was similar in concept to the system eventually

installed on the space shuttle orbiters. The ill-fated flight 2-45-81 marked the first use of the retractable pitot tube, in parallel with the normal system. Despite other problems on the flight,

Bob Rushworth considered the new system acceptable for flight, and subsequent data analysis confirmed this.[318]