When the Langley Aeronautical Laboratory was transferred to NASA it was renamed the Langley Research Center. The name change and switch from NACA to NASA signified a change in emphasis rather than a radically new mission. Indeed, Langley had been researching high-speed flight, developing heatresistant materials and structures, and firing multistage research rockets from Wallops Island, Virginia, for almost a decade. With its several hypersonic wind tunnels and wide experience with rocket testing, Langley became a cornerstone of NASA's space effort. With a mandate to place man into space and on the Moon quickly, NASA drew heavily on Langley's expertise and facilities. The Space Task Group, which led the Mercury, Gemini, and Apollo programs, was staffed mainly from Langley personnel.
There was already in place at Langley a wide spectrum of hypersonic/hypervelocity facilities. To meet the new challenges of manned space flight and NASA's aeronautical assignment, three important new wind tunnels were built during the 1960s.
The Langley Continuous Flow Hypersonic Tunnel, under the guidance of Eugene Love, entered the planning stage in 1958, essentially the same time NASA was formed. Preceding hypersonic tunnels had been blowdown facilities. Runs were short-just a few minutes at most-and data productivity and experimental versatility were wanting. The goals for the new tunnel were ambitious: continuous operation at Mach numbers up to 12.
The first major problem facing the tunnel engineers was the very high pressure ratio (about 700) that had to be maintained continuously for Mach 12 operation. Initially, six big air compressors were to be lined up in series, the output of the first feeding the input of the second, and so on down the line. The highpressure air would then pass through a square nozzle, 1-1/4 inches on a side at the throat, and into a test section 31 inches square. Discharge was to be into a vacuum sphere. The system worked on paper, but the sixth and final compressor in the chain pushed the state of the art too far. It had to be dropped, and the design goal was compromised to Mach 10.
Heat was the second major worry. It had to be added in the settling chamber before the nozzle to prevent air liquefaction in the nozzle and then extracted ahead of the vacuum sphere to maintain the...
.... sphere's structural integrity and protect the downstream compressors. Adding heat was relatively simple: a 13 000-kilowatt electric resistance heater in the settling chamber raised the air temperature to 1450° F. The air temperature dropped rapidly in the nozzle as heat energy was converted to kinetic energy. But in the vacuum sphere, the process reversed as the air slowed down. The kinetic energy was transformed back into heat. A large water cooler had to be installed to pull the temperature of the air in the vacuum sphere down to about 100° F. Another hot spot was located in the nozzle throat where the air was dense, moving at Mach 1, and still about 1450° F. Distilled water circulating within the nozzle walls kept temperatures within bounds.
The continuous flow hypersonic tunnel went on line in 1962 with blowdown capabilities only. Two years later the addition of a compressor system converted it to continuous operation. Through the years, this facility has been applied primarily to the study of the aerodynamic performance and heat transfer on winged reentry vehicles, such as the Space Shuttle.
Officially known as the 8-foot High-Temperature Structures Tunnel, the goal of this facility was the realistic testing of flight structures under the stresses and high temperatures of hypersonic flight. In charge of this effort was Langley's Robert Howell. Existing hypersonic tunnels, even though capable of continuous flow, could not duplicate the structural problems encountered at these high velocities. Test section size was the basic limitation. Small-scale models were adequate for aerodynamic testing, but the buildup of thermal stresses in complex aerospace vehicles could...
 ....best be studied under full-scale conditions where internal structures could be duplicated.
The Langley 9 x 6-foot thermal structures tunnel which went on line in 1956, was the right size but it only reached Mach 3; Mach 7 was the target of the new 8-foot tunnel. At this airspeed, the energy requirement was prodigious: 1 000 000 horsepower. The electrical equivalent (746000 kilowatts) represented the full capacity of a very large commercial electrical generating station. Electrical power plants in the Langley region could not divert such a large block of power even for a few minutes; a methane blowtorch was offered as a solution. By burning methane in air at very high pressures and expanding the combustion products through a hypersonic nozzle, Mach 7 could be attained in an 8-foot test section. But would the combustion products of methane (mostly carbon dioxide and water vapor) simulate air closely enough? Analysis showed that the flow parameters would deviate less than 10 percent. The methane torch was worth a try.
The heart- of the new tunnel was the methane burner, which required the combustion of 1000 pounds of methane gas per second at 270 atmospheres pressure and a temperature of 3500° F. These conditions were well beyond the state of the art in the late....
....1950s, but 200 atmospheres at 3000° F seemed attainable. The construction contracts were awarded in 1960. Cold test runs of the completed tunnel occurred in 1964, but high-temperature runs had to wait until 1968.
The high temperature structures tunnel, because of the copious combustion products, had to be of the nonreturn type; that is, the gases are not recirculated. A huge tank farm and methane storage complex feed fuel to the burner. The flow is through the nozzle, then through the test section, past an annular injector that lowers the exit pressure during startup, and finally out through a diffuser into a swampy area. Portions of flight vehicles to be tested can be inserted into the gas stream as quickly as one second, and withdrawn in the same period. Pieces of failed test structures simply fly out the aft end of the tunnel into the uninhabited swamp.
This unique tunnel came too late to be useful in the Apollo Program, but it has been of immense benefit in testing the Space Shuttle and hypersonic aircraft. In fact, it was almost as if the tunnel were designed specifically for the Space Shuttle. Its test conditions have been ideal for full-scale testing of the insulating tiles that preserve the integrity of the Space Shuttle during reentry. It represents still another case of serendipity, for the high temperature structures tunnel was conceived almost 20 years before the Space Shuttle was built.
As explained earlier in this chapter, an important fraction of the heat that eats away at the nose of a reentering spacecraft radiates from the incandescent cap of gas that piles up between the nose of the vehicle and the nose shock. Ames Research Center developed several shock tubes and other high-velocity devices to study the high-temperature gas properties in this critical region. Langley aerodynamicists, in the never-ending search for effective hypervelocity simulation, introduced the Hot-Gas Radiation Research Laboratory in 1969. The core facilities included a high-performance, 6-inch-diameter arc-driven shock tube and a 6-inch-diameter expansion tube. These facilities are driven by the discharge of a 10-megajoule energy storage bank. The expansion tube consisted of three stages: (1) a driver section typically filled with helium at 350 atmospheres (for unheated operation), (2) a driven section containing the test gas at about 0.05 atmosphere, and (3) an acceleration section at...
...about 0.0005 atmosphere. After rupture of the diaphragm in the driver section, the pressure ratio across the three stages builds up to several million-a ratio sufficient to reach Mach 15 and flow velocities to 25 000 feet per second. Separating the second and third stages is a gossamer-like plastic diaphragm about one-tenth the thickness of this page. Its strength is just sufficient to withstand the initial pressure difference between the second and third stages. Upon rupture by the shock wave in the driven section, the attendant expansion doubles the velocity of the gas entering the expansion section.
The pulse of test gas impacting the model lasts only about 300 microseconds, but it is smooth and possesses a very low turbulence level. This is followed by the impact of the high-pressure driver gas which strikes the model with a sledgehammer-like blow. Instrument response must be faster than a microsecond to provide meaningful test data. Furthermore, a boundary layer builds up quickly around the walls of the test section so that only the central core of gas- about 3 inches in diameter-is really useful. Despite these time and space limitations, surprisingly good schlieren photos, pressure measurements, heat transfer information, and other flow data can be recorded in the fraction of a thousandth of a second of useful run time.