Materials Research and Development: The NASP Legacy

The National Aero-Space Plane (NASP) program had much to contribute to metallurgy, with titanium being a particular point. It has lately come to the fore in aircraft construction because of its high strength-to-den – sity ratio, high corrosion resistance, and ability to withstand moderately
high temperatures without creeping. Careful redesign must be accom­plished to include it, and it appears only in limited quantities in aircraft that are out of production. But newer aircraft have made increasing use of it, including the two largest manufacturers of medium – and long – range commercial jetliners, Boeing and Airbus, whose aircraft and their weight of titanium is shown in Table 4.

TABLE 4: BOEING AND AIRBUS AIRCRAFT MAKING SIGNIFICANT USE OF TITANIUM
AIRCRAFT (INCLUDING THEIR ENGINES)	WEIGHT OF TI, IN METRIC TONS
Boeing 787	134
Boeing 777	59
Boeing 747	45
Boeing 737	18
Airbus A380	145
Airbus A350	74
Airbus A340	32
Airbus A330	18
Airbus A320	12

These numbers offer ample evidence of the increasing prevalence of titanium as a mainstream (and hence no longer "exotic”) aviation mate­rial, mirroring its use in other aspects of the commercial sector. For example, in the 1970s, the Parker Pen Company used titanium in its T-1 line of ball pens and rollerballs, which it introduced in 1971. Production stopped in 1972 because of the high cost of the metal. But hammerheads fabricated of titanium entered service in 1999. Their light weight allows a longer handle, which increases the speed of the head and delivers more energy to the nail while decreasing arm fatigue. Titanium also substan­tially diminishes the shock transferred to the user because it generates much less recoil than a steel hammerhead.

In advancing titanium’s use, techniques of powder metallurgy have been at the forefront. These methods give direct control of the micro­structure of metals by forming them from powder, with the grains of powder sintering or welding together by being pressed in a mold at high temperature. A manufacturer can control the grain size independently of any heat-treating process. Powder metallurgy also overcomes restrictions on alloying by mixing in the desired additives as powdered ingredients.

Several techniques exist to produce the powders. Grinding a metal slab to sawdust is the simplest, though it yields relatively coarse grains. "Splat-cooling” gives better control. It extrudes molten metal onto the chilled rim of a rotating wheel that cools it instantly into a thin ribbon. This represents a quenching process that produces a fine-grained micro­structure in the metal. The ribbon then is chemically treated with hydro­gen, which makes it brittle so that it can be ground into a fine powder. Heating the powder then drives off the hydrogen.

The Plasma Rotating Electrode Process, developed by the firm of Nuclear Metals, has shown particular promise. The parent metal is shaped into a cylinder that rotates at up to 30,000 revolutions per min­ute (rpm) and serves as an electrode. An electric arc melts the spinning metal, which throws off droplets within an atmosphere of cool inert helium. The droplets plummet in temperature by thousands of degrees within milliseconds and their microstructures are so fine as to approach an amorphous state. Their molecules do not form crystals, even tiny ones, but arrange themselves in formless patterns. This process, called "rapid solidi­fication,” has brought particular gains in high-temperature strength.

Standard titanium alloys lose strength at temperatures above 700 to 900 °F. By using rapid solidification, McDonnell-Douglas raised this limit to 1,100 °F prior to 1986, when NASP got underway. Philip Parrish, the manager of powder metallurgy at the Defense Advanced Research Projects Agency (DARPA), notes that his agency spent some $30 mil­lion on rapid-solidification technology in the decade after 1975. In 1986, he described it as "an established technology. This technology now can stand along such traditional methods as ingot casting or drop forging.”[1095]

Eleven-hundred degrees nevertheless was not enough. But after 1990, the advent of new baseline configurations for the X-30 led to an appre­ciation that the pertinent areas of the vehicle would face temperatures no higher than 1,500 °F. At that temperature, advanced titanium alloys could serve in metal matrix composites (MMCs), with thin-gauge met­als being reinforced with fibers.

A particular composition came from the firm of Titanium Metals and was designated Beta-21S. That company developed it specifi­cally for the X-30 and patented it in 1989. It consisted of Ti along with 15Mo+2.8Cb+3Al+0.2Si. Resistance to oxidation proved to be its strong suit, with this alloy showing resistance that was two orders of magnitude greater than that of conventional aircraft titanium. Tests showed that it could also be exposed repeatedly to leaks of gaseous hydrogen without being subject to embrittlement. Moreover, it lent itself readily to being rolled to foil-gauge thicknesses of 4 to 5 mils in the fabrication of MMCs.[1096]

There also was interest in using carbon-carbon for primary struc­ture. Here the property that counted was not its heat resistance but its light weight. In an important experiment, the firm of LTV fabricated half of an entire wing box of this material. An airplane’s wing box is a major element of aircraft structure that joins the wings and provides a solid base for attachment of the fuselage fore and aft. Indeed, one could compare it with the keel of a ship. It extends to left and right of the air­craft centerline, with LTV’s box constituting the portion to the left of this line. Built at full scale, it represented a hot-structure wing proposed by General Dynamics. It measured 5 by 8 feet, with a maximum thickness of 16 inches. Three spars ran along its length, five ribs were mounted transversely, and the complete assembly weighed 802 pounds.

The test plan called for it to be pulled upward at the tip to repro­duce the bending loads of a wing in flight. Torsion or twisting was to be applied by pulling more strongly on the front or rear spar. The maxi­mum load corresponded to having the X-30 execute a pullup maneuver at Mach 2.2 with the wing box at room temperature. With the ascent continuing and the vehicle undergoing aerodynamic heating, the next key event brought the maximum difference in the temperatures of the top and bottom of the wing box, with the former being at 994 °F and the latter being at 1,671 °F. At that moment, the load on the wing box corre­sponded to 34 percent of the Mach 2.2 maximum. Farther in the flight the wing box was to reach peak temperature, 1,925 °F, on the lower surface. These three points were to be reproduced through mechanical forces applied at the ends of the spars and through the use of graphite heaters.

But several important parts delaminated during their fabrication, which seriously compromised the ability of the wing box to bear its specified loads. Plans to impose the peak or Mach 2.2 load were aban­doned, with the maximum planned load being reduced to the 34 percent associated with the maximum temperature difference. For the same rea­son, the application of torsion was deleted from the test program. Amid these reductions in the scope of the structural tests, two exercises went forward during December 1991. The first took place at room tempera­ture and successfully reached the mark of 34 percent without causing further damage to the wing box.

The second test, a week later, reproduced the condition of peak tem­perature difference while briefly applying the calculated load of 34 per­cent. The plan then called for further heating to the peak temperature of 1,925 °F. As the wing box approached this value, a difficulty arose because of the use of metal fasteners in its assembly. Some were made from coated columbium and were rated for 2,300 °F, but most were a nickel alloy that had a permissible temperature of 2,000 °F. However, an instrumented nickel-alloy fastener overheated and reached 2,147 °F. The wing box showed a maximum temperature of 1,917 °F at that moment, and the test was terminated because the strength of the fasteners now was in question. This test nevertheless counted as a success because it had come within 8 degrees of the specified temperature.[1097]

Both tests thus were marked as having achieved their goals, but their merits were largely in the mind of the beholder. The entire project would have been far more impressive if it had avoided delamination, had successfully achieved the Mach 2.2 peak load, and had subjected the wing box to repeated cycles of bending, torsion, and heating. This effort stood as a bold leap toward a future in which carbon-carbon might take its place as a mainstream material, but it was clear that this future would not arrive during the NASP program. However, the all-carbon – composite airplane, as distinct from one of carbon-carbon, has now become a reality. Carbon alone has high temperature resistance, whereas carbon composite burns or melts readily. The airplane that showcases carbon composites is the White Knight 2, built by Burt Rutan’s Scaled Composites firm as part of the Virgin Galactic venture that is to achieve commercial space flight. As of this writing, White Knight 2 is the world’s largest all-carbon-composite aircraft in service; even its control wires are carbon composite. Its 140-foot-span wing is the longest single car­bon composite aviation component ever fabricated.[1098]

Far below this rarefied world of transatmospheric tourism, carbon composites are becoming the standard material for commercial avia­tion. Aluminum has held this role up to Mach 2 since 1930, but after 80 years, Boeing is challenging this practice with its 787 airliner. By weight, the 787 is 50 percent composite, 20 percent aluminum, 15 per­cent titanium, 10 percent steel, and 5 percent other. The 787 is 80 per­cent composite by volume. Each of them contains 35 tons of composite reinforced with 23 tons of carbon fiber. Composites are used on fuse­lage, wings, tail, doors, and interior. Aluminum appears at wing and tail leading edges, with titanium used mainly on engines. The exten­sive application of composites promotes light weight and long range. The 787 can fly nonstop from New York City to Beijing. The makeup of the 787 contrasts notably with that of the Boeing 777. Itself consid­ered revolutionary when it entered service in 1995, it nevertheless had a structure that was 50 percent aluminum and 12 percent composite. The course to all-composite construction is clear, and if the path is not yet trodden, nevertheless, the goal is clearly in sight. As in 1930, when all-metal structures first predominated in American commercial avia­tion, the year 2010 marks the same point for the evolution of commer­cial composite aircraft.

The NASP program also dealt with beryllium. This metal had only two-thirds the density of aluminum and possessed good strength, but its temperature range was restricted. The conventional metal had a limit of some 850 °F, while an alloy from Lockheed called Lockalloy, which contained 38 percent aluminum, was rated only for 600 °F. It had never become a mainstream material like titanium, but, for the X-30, it offered the advantage of high thermal conductivity. Work with titanium had greatly increased its temperatures of use, and there was hope of achiev­ing similar results with beryllium.

Initial efforts used rapid-solidification techniques and sought tem­perature limits as high as 1,500 °F. These attempts bore no fruit, and from 1988 onward the temperature goal fell lower and lower. In May 1990, a program review shifted the emphasis away from high-tempera­ture formulations toward the development of beryllium as a metal suit­able for use at cryogenic temperatures. Standard forms of this metal became unacceptably brittle when only slightly colder than -100 °F, but cryoberyllium proved to be out of reach as well. By 1992, investiga­tors were working with ductile alloys of beryllium and were sacrificing all prospect of use at temperatures beyond a few hundred degrees but were winning only modest improvements in low-temperature capabil­ity. Terence Ronald, the NASP materials director, wrote in 1995 of rapid – solidification versions with temperature limits as low as 500 °F, which was not what the X-30 needed to reach orbit.[1099]

In sum, the NASP materials effort scored a major advance with Beta – 21S, but the genuinely radical possibilities failed to emerge. These included carbon-carbon as primary structure along with alloys of beryllium that were rated for temperatures well above 1,000 °F. The latter, if available, might have led to a primary structure with the strength and temperature resistance of Beta-21S but with less than half the weight. Indeed, such weight savings would have ramified throughout the entire design, lead­ing to a configuration that would have been smaller and lighter overall.

Generally, work with materials fell well short of its goals. In dealing with structures and materials, the contractors and the National Program Office established 19 program milestones that were to be accomplished by September 1993. A General Accounting Office program review, issued in December 1992, noted that only six of them would indeed be com­pleted.[1100] This slow progress encouraged conservatism in drawing up the bill of materials, but this conservatism carried a penalty.

When the scramjets faltered in their calculated performance and the X-30 gained weight while falling short of orbit, designers lacked recourse to new and very light materials, such as beryllium and carbon-carbon, that might have saved the situation. With this, NASP spi­raled to its end. The future belonged to other less ambitious but more attainable programs, such as the X-43 and X-51. They, too, would press the frontier of aerothermodynamic structural design, as they pioneered the hypersonic frontier.