HYPERSONIC FACILITIES COMPLEX (HFC) Significance

The Hypersonic Facilities Complex, originally named the Gas Dynamics Laboratory,

was the NACA’s earliest installation with multiple hypersonic wind tunnels. The various

tunnels utilized air and, later, specific gasses, particularly helium and nitrogen, to investigate

aerodynamic and aerothermodynamic problems in flight regimes above the speed range of

Mach 5. The facility’s core tunnels have been blowdown-type tunnels powered by common

differential-pressure systems.

Designed to be flexible and adaptable, the Hypersonic Facilities Complex has seen

numerous modifications and managerial reorganizations since its opening in 1951, and it

remains one of NASA’s premier resources for very-high-speed and spacecraft re-entry

studies.

Description

The Hypersonic Facilities Complex (HFC), Building 1247 in Langley’s West Area,

is a U-shaped structure of brick and concrete. The central portion of its southeast face is a 3-

story brick office building with a concrete portal surrounding its central entrance. This

central portion is flanked by two concrete wings extending in a northwesterly direction. The

“east wing” (Building 1247D) contains rooms and test cells housing most of the laboratory’s

tunnels that use air. The “west wing” (Building 1247B) historically housed the helium and

nitrogen tunnels, although some air-based tunnels are located in this wing as well. Both

wings feature central halls with test cells on each side, though their plans differ in size and

proportion. The East Wing is the larger of the two, and it features a wider central hall. A

partial second floor in the East Wing contains a central control room that manages air

distribution to all blowdown tunnels in the wing.

Reinforced concrete construction was used in the two wings of test cells as a

precaution against the very high pressures and temperatures required for hypersonic

blowdown tunnels. Heavy steel doors with equally heavy hardware open from each test cell

onto the central hall. Each test cell has floor and wall penetrations for air/gas supply and

vacuum pipes to suit the requirements of its specific tunnel.

Visually, the HFC is dominated by five large, steel vacuum spheres, four 60-foot-

diameter spheres along the southwest side, and a 100-foot-diameter, ribbed one between the

two wings. These are supplemented by several other, smaller vacuum spheres between the

wings. Large-diameter steel pipes, which are exposed to the extent practical, connect these

spheres to the tunnels and vacuum pumps inside the structure. Through the use of large

valves, many of the tunnel–sphere connection configurations can be changed to suit the

nature and required run time of each test.

Room B-104 in the West Wing contains equipment for handling and reclaiming

helium. Following a run in any of the helium tunnels, this equipment pumps the helium

from the vacuum sphere(s), filters it, and compresses it to 5,000 pounds per square inch for

storage in an on-site rack of high-pressure bottles. From there, helium could be routed to

any of the helium tunnels when needed, though only one could be run at a time.

The open area partially surrounded by the building and extending toward the

northwest contains three of the large vacuum spheres, a cooling tower, racks of high-

pressure gas bottles, and vacuum pumps for the air system. Water circulated through the

cooling tower is used to cool the hot air or gas leaving many of the tunnels and, thus,

improve the efficiency of the vacuum spheres.

History

The Hypersonic Facilities Complex (HFC) was part of the large investment

Congress made in NACA facilities during the late 1940s and early 1950s to support high-

speed-flight research. Wind tunnels for the supersonic and transonic flight regimes were

conventional, closed-circuit tunnels. Large, powerful, and expensive, the new tunnels,

particularly the Unitary Plan Wind Tunnels, for speeds up to about Mach 4.5 generally

resembled their subsonic predecessors, even though their test sections incorporated key new

features and their drive power increased dramatically. Wind tunnels for speeds in the

hypersonic range (Mach 5 and higher) presented conditions too extreme to achieve with this

design. As in supersonic tunnels, converging-diverging nozzles can be designed to produce

hypersonic speeds, but the high pressure and temperature differentials involved mean that

the chemical and physical properties of the test gas must be considered. Hypersonic nozzles

typically have exit-to- throat ratios on the order of 20 – 30 to 1. Most gasses, including air,

will liquefy leaving the throat of such nozzles, thanks to a drastic drop in temperature. To

prevent liquefaction, the gas must be heated to 1,000 F, or more, before it enters the nozzle.

This means that most nozzles need to be cooled to prevent distortion and degradation.

Similarly, high pressures, often above 1,000 pounds per square inch, are necessary to

achieve the proper mass flow through the nozzle. Such extremes dictated a facility unlike

any other at Langley.1

Langley’s 11-Inch hypersonic Tunnel, opened in 1947, had shown that much could

be learned from short test runs, but it also demonstrated how hypersonic tunnels had to be

carefully designed for a given speed. The nozzle requirements, and their concurrent

pressure and temperature demands, were too exacting to allow the use of flexible walls to

vary test speeds. While it was theoretically possible to design a hypersonic tunnel with

sufficient compressors to achieve continuous operation, the power and expense made that an

impractical choice, so engineers decided to pursue blowdown-type tunnels. These could use

reasonably sized compressors to compress air, or any other gas, over time and, similarly,

vacuum pumps could evacuate vacuum vessels in advance of test runs. With quick-acting

valves, rapid-responding instrumentation, and careful management, two, or more, test runs

might be possible from a single preparation cycle.

John Stack, the indomitable head of Langley’s Compressibility Research Division,

championed the idea of a facility that had a common gas system with a large pressurized

sphere serving several blowdown tunnels. Each tunnel would be designed to different

specifications so that they complimented one another, giving the facility a wide range of

testing capabilities. Antonio Ferri, an Italian researcher who had joined his staff after World

War II, had experience with high-speed blowdown facilities in Europe, agreed in principle,

but Ferri and engineer Macon C. Ellis, took the idea a step further. They replaced the single,

large pressure sphere with a rack of small-diameter bottles capable of holding air at a much-

higher pressure, as much as 5,000 pounds per square inch (borrowing a technique that had

worked well aboard submarines). They then converted Stack’s big sphere into a vacuum

vessel. Connecting tunnels between the two would yield the highest possible differential

pressure across the tunnels’ nozzles.2

This multi-tunnel concept for the Gas Dynamics Laboratory received approval from

Congress in 1949. At $5.5 million, its initial cost was more than twice that of any other

facility at Langley, though the price tag for Langley’s Unitary Plan Tunnel, approved later

the same year, would triple that staggering figure. During these early years of the Cold

War—the Soviet Union detonated its first atomic bomb in 1949—Congress was in the mood

to maximize America’s investment in aeronautical research and, thus, ensure its strength in

high-speed aviation and the emerging field of guided missiles.

At its opening in 1951, the Gas Dynamics Laboratory housed several supersonic and

hypersonic blowdown tunnels using air and helium to cover speeds ranging from Mach 1.5

to Mach 18. Ferri, and Ellis, with design assistance from Clinton Brown, developed their

ideas into a complex, but flexible and practical distribution system for dried air and helium

that serviced test cells arranged in two wings around a central office building. The test

sections for these tunnels all measured about 20 inches, a size that allowed reasonably sized

models and nozzle dimensions that were practical to machine and operate. A bottle rack

held 19,000 cubic feet of air at 5,000 pounds per square inch. By routing this air through

one of the tunnels and then to a vacuum sphere, pressure ratios across the nozzle as high as

10,000 to 1 were possible. Steam and electric heaters heated the supply air to temperatures

approaching 1100 Fahrenheit to prevent liquefaction, and a cooling system furnished water

to cool nozzles and the tunnel exhaust flow. Steam ejectors lowered the pressure in the

vacuum spheres. A similar system handled helium, but it included vacuum pumps and

purification equipment and a compressor to recover the helium for re-use. Given the

unknowns and extreme conditions of hypersonic research, the engineers specified

reinforced-concrete test cells to enclose the various tunnels and prevent an accident in one

from affecting the others. This basic design has been maintained to this day in spite of

numerous changes to the facility’s tunnels.3

Many of the problems tackled by HFC researchers in its early years involved

conditions that would be experienced by guided missiles, especially by nuclear warheads re-

entering the atmosphere on the downward side of their ballistic trajectories. At Mach 5 –

10, the problems involved not only aerodynamic stability, but also searing heat generated by

both air friction in the boundary layer and the high compression in shock waves. Finding

designs that minimized the heat exposure while maintaining a predictable flight path and

materials that could withstand 2,000-plus Fahrenheit temperatures quickly became inter-

related problems that NACA researchers tackled at both Ames and Langley.4

In the midst of this boom in research, the preferred method of hypersonic research

became something of a controversy within the NACA. The Pilotless Aircraft Research

Division (PARD) had been established at Wallops Island, Virginia, during World War II as

a site for launching small rockets that could boost model airplanes to speeds and altitudes in

free flight that could not then be duplicated in Langley’s wind tunnels. By the time the Gas

Dynamics Laboratory opened, Langley had closed the gap considerably, but researchers at

Wallops argued that the technology they had developed produced superior data, especially

as speeds approached Mach 10. The NACA appointed an ad-hoc committee consisting of

Clinton Brown from Compressibility research, Charles Zimmerman from Stability and

Control, and William O’Sullivan from PARD to investigate and make recommendations.

Their findings were mixed. While acknowledging that PARD’s methods were better at

generating the high temperatures desired, Langley’s (and Ames’) hypersonic tunnels could

do most of the research and development work at a lower cost. Flight testing was necessary,

but more cost-effective for final testing of developed hardware.5

Blowdown tunnels offered something else for researchers that PARD flights could

not provide: the ability to test with gasses other than air. Helium offered two advantages.

The speed of sound (Mach 1) in helium is almost three times that in air at the same

temperature. Secondly, helium does not liquefy until its temperature reaches -452

Fahrenheit, only a few degrees above absolute zero, making heating unnecessary up to about

Mach 25, and reasonable heat permits speeds as high as Mach 40, beyond what PARD

flights could achieve. With the proper equipment, helium could be highly purified and re-

used, thus eliminating problems stemming from water and most other contaminants.

Helium did, however, have some disadvantages. At elevated temperatures, helium, a

monoatomic gas, had different thermodynamic properties than air, which is composed

mostly of diatomic molecules, and the required corrections were not well understood in the

early 1950s. Langley researchers have also used other gasses, including nitrogen and

tetrafluoromethane (CF4), in some of the HFC’s tunnels to reduce operating problems and

improve their understanding of gas behavior under extreme conditions.

Research operations and equipment assignments in the Hypersonic Facilities

Complex have changed many times since its opening, and its tunnels have been removed,

installed, modified, and relocated in a bewildering manner, but it is the intelligent initial

design of such a flexible facility that has made these alterations to meet changing needs

possible. Compared to the cost of new construction that might otherwise be needed to

accommodate new tunnels, the savings to NASA, though difficult to estimate, have been,

and continue to be, enormous.

Tunnels in the Hypersonic Facilities Complex

The Hypersonic Facilities Complex, intentionally designed for flexibility, has housed

numerous wind tunnels during its life. Some of these remain in service, while others have

been dismantled or modified. Not all of them are hypersonic tunnels, since some space

often has been available for other research, and slower tunnels have frequently occupied

some rooms. In certain cases, the information available about a particular tunnel is

imprecise or unavailable, but the primary tunnels that have been, or remain, in the

Hypersonic Facilities Complex are briefly described below. For clarity, these will be

presented in room-number order rather than chronological order. Only rooms (test cells)

with aerodynamic tunnels are shown. Facilities intended primarily for studies in fields such

as magnetoplasmadynamics, materials, and acoustics are not included. The HFC also has

hosted a number of small, “table-top” hypersonic tunnels over the years, but records about

them are scarce and these devices are not included.

In the context of hypersonic tunnels, a “conventional” tunnel is a blowdown tunnel,

whether the inlet gas is heated or unheated, that is capable of test runs at least several

seconds long. “Shock,” “impulse,” “expansion tube,” and “hotshot” all refer to facilities that

utilize sudden, essentially explosive driving forces to produce ultra-high-speed test flows

lasting only fractions of a second.

Many of these tunnels have had multiple names, both official and unofficial, over the

years, with no uniform naming rationale apparent. The primary names shown here are those

that appear to be the most consistent and recent, with the known alternate names indicated in

the text.6

B-102 – 22-Inch Helium Tunnel

This conventional tunnel, opened in 1960, was originally named the Hypersonic

Helium Tunnel. Three axisymmetric nozzles for it were built to deliver Mach 18, 22, and 26

flows, but only the Mach 22 nozzle has been calibrated, and it has been the one most-

frequently utilized. The test section measures 221/2 inches in diameter, which provides a

core test region between 8 and 10 inches in diameter. Helium is supplied from high-

pressure bottles and exhausted into one, or more, vacuum spheres. A hydraulically-actuated

mechanism inserts a model into the stabilized flow, and a second, similar device does the

same for a shield that protects the model from start-up transients or debris. The flow is

established by opening quick-acting valves and stabilized with a quick-response nozzle-

throat plug. Helium is recovered from the vacuum sphere after a test for re-use. This

tunnel’s primary modification involved the addition of a second circuit featuring a larger test

chamber known as the Open-Jet Leg. The original circuit, then referred to as the

Aerodynamic Leg, remained in service until being closed about 2003. It was demolished

about 2007.7

B-105 – Hypersonic Helium Tunnel Open-Jet Leg

The Open-Jet Leg was an addition to the 22-Inch Helium Tunnel that opened in

1969. It utilized two axisymmetric nozzles, on that produced Mach 20 flow at its 22-inch

exit and a second that generated Mach 40 flow at its 36-inch exit. Either discharged into a

larger test section, where a model-injection device swung the model into the flow centerline

after steady flow had been established. The support mechanism allowed the angle of attack

to vary ±20 degrees. As with the Aerodynamic Leg, heat was not required to prevent

liquefaction during Mach 20 tests, but heat from the building’s heaters was needed for those

at Mach 40. Run time was approximately 20 seconds. With a declining interest in

hypersonic research and budgetary constraints during the 1990s, NASA decided to close this

tunnel and subsequently demolish it tunnel about 1998.8

B-107 – 60-Inch Mach 18 Helium Tunnel

One of the facility’s original tunnels, this conventional tunnel, also known as the

Mach 20 High-Reynolds Number Helium Tunnel and as one of the Hypersonic

Aeroelasticity Tunnels, opened in 1952 as the NACA’s largest helium tunnel. It featured

axisymmetric nozzles for Mach 16.5 and Mach 18 using gas supplied by the building’s

5,000 pound-per-square-inch helium system and recovered from the 60-foot-diameter

vacuum spheres on the building’s southwest side. Most runs involved the use of two

spheres and lasted 5 – 10 seconds. Models were sting- or strut-mounted with a variable

angle-of-attack adjustment. An electron-bombardment system for flow visualization was

available, as was a spark schlieren photography system. Ambient temperatures were

adequate to avoid helium liquefaction for tests in this Mach range. This 60-inch tunnel

shared the test chamber and a control room with a similar 37-inch tunnel. Much of the

research performed here, especially during its later years, involved turbulent boundary

layers. With a declining interest in hypersonic research and budgetary constraints during the

1990s, NASA decided to demolish this tunnel in 1998.9

37-Inch Mach 10 Helium Tunnel

Also known as the Mach 10 High-Reynolds Number Helium Tunnel and as one of

the Hypersonic Aeroelasticity Tunnels, this tunnel opened in 1952. It was essentially a

smaller, slower version of its 60-inch sibling, and it operated in the same manner using the

same helium system. The two tunnel circuits shared a common control room. This tunnel

ceased operations in the late 1970s, and it was demolished along with the 60-inch circuit in

1998.10

2- by 3-Foot Low-Boundary-Layer Tunnel

This atmospheric, closed-circuit tunnel was installed in 19__ as part of a

modification to room H-107 in Building 1247’s west wing. (Room B-107 had been re-

designated H-107 during a prior modification to the space.) When originally installed in this

room, it was oriented in a north-south direction, and it shared the space with the 60- and 37-

inch helium tunnels until those tunnels were demolished. Langley considered relocating it

to an extension of Building 1214, which housed the Basic Aerodynamics Research Tunnel

(BART) in 1997, but rejected that option, even though the two tunnels performed similar

work. Instead, the tunnel underwent a major rehabilitation during 2004 that rotated it 90

degrees to better fit in the H-107 space.

The tunnel is a self-contained device capable of speeds up to approximately 100

miles per hour, and it incorporates a honeycomb, four screens, and a 10:1 contraction ratio

to reduce turbulence. The floor and ceiling of the 20-foot-long test section are adjustable to

allow operators to set a desired longitudinal pressure gradient. Additionally, the tunnel is

equipped with acoustic drivers upstream and downstream of the test section to excite the

flow as needed.11

B-109 – 20-Inch Mach 17 Nitrogen Tunnel

Previously known as the Hypersonic Nitrogen Tunnel, this conventional tunnel was

built in 1964. Nitrogen gas, flashed from a liquid storage tank, passed through an electric

resistance heater at the nozzle’s inlet, continued through the tunnel’s axisymmetric nozzle

and test section, and exited it through a diffuser and aftercooler into a vacuum sphere.

Because of the benign nature of nitrogen and its low cost, the gas was vented to atmosphere

after test runs and not recovered. Despite its name, the tunnel’s test section was shown as

being only 17 inches in diameter in Langley facility résumés, though it may originally have

had a larger test chamber and on open jet. This nitrogen tunnel was built to investigate how

well a single gas could simulate air at hypersonic flight speeds. Unfortunately, the flow in

this tunnel was more turbulent than desired, and nitrogen offered no clear advantages over

air that justified its cost. Proposals to rectify the flow problems were not approved. With a

declining interest in hypersonic research and budgetary constraints during the 1990s, NASA

decided to close this tunnel about 1998 and subsequently demolish it in 2007.12

B-111 – Arc-Heated Scramjet Test Facility

Originally known as the 4-Foot Hypersonic Arc tunnel, this tunnel was built in 1964

to support high-enthalpy hypersonic fluid mechanics research. (Enthalpy is a

thermodynamic term for the total internal energy of a volume of gas, including its

temperature and pressure, expressed as heat. Its English measurement unit is Btu/lbm.) The

tunnel used an electric arc with a 10 to 20 megawatt direct-current power supply to heat dry

inlet air from the central utilities system. The heated air then passed through an

axisymmetric nozzle and accelerated into either a 2- or 4-foot-diameter test section (the core

test diameter was about one-half that) and impinged on the model. The air then exhausted

through an aftercooler and into the 100-foot-diameter vacuum sphere. Since a different

nozzle length was needed for each test section, the heater and nozzle were mounted on a

track that guided the necessary movement. It could generate speeds between Mach 8 and

18, with enthalpies between 1,500 and 6,000 Btu/lb.

This tunnel was converted into a facility for testing complete scramjet models in

1974 and re-named the Scramjet Test Facility, later refined into Arc-Heated Scramjet Test

Facility (AHSTF), to distinguish it from combustion-heated scramjet tunnels in Building

1221. The conversion involved changes to the nozzle, model support hardware, and

instrumentation, as well as the addition of a model fuel system, but its basic, arc-heated

operation remained the same. Managed by the Hypersonic Air-breathing Propulsion

Branch, it remains in active service, testing scramjet models approximately 8 by 10 inches in

cross section at simulated flight speeds of Mach 4.7 – 8.0. Inlet air from a 5,000-pound-per-

square-inch bottle field passes through an electric arc, where it is heated to as high as 4,600

Fahrenheit, accelerated through the nozzle into the test section, and, as before, exhausted

through an aftercooler to the 100-foot-diameter vacuum sphere. Engine fuel is gaseous

hydrogen, with a 20/80-percent silane/hydrogen mixture available to promote ignition. Run

times are typically 10 – 60 seconds. During 2008, the AHSTF was reconfigured to run in a

direct-connect mode, which simulates inlet conditions at the engine’s combustor for tests

involving fuel ignition and flame stability.13

B-115 – Hotshot Tunnel

The Hotshot Tunnel, built in 1960 and operational in November 1962, was an arc-

heated, impulse tunnel. Being Langley’s first hotshot tunnel, setup and calibration proved

arduous, with difficult problems to be solved in materials electrodes, insulators, connections,

and instrumentation. It was built primarily to perform force, pressure-distribution, and heat-

transfer tests at Mach 20 and above.

As with all hotshot tunnels, this tunnel operated by heating its test gas (air, nitrogen,

or helium) in a small chamber at one end with an electric arc. A bank of 720 capacitors

delivered an instantaneous 7,500-volt, 2 x 106-joule potential across two electrodes in the

chamber, and the arc created heated the gas to about 5,000 Fahrenheit, which, in turn,

caused the gas pressure to increase to approximately 10,000 pounds per square inch. This

pressure burst a polyethylene terephthalate diaphragm and allowed the gas to accelerate

through an asymmetric nozzle and pass through the test section into a 200-cubic-foot dump

tank. Prior to the run, the entire tunnel downstream of the diaphragm had been evacuated to

approximately 1 micron of mercury by mechanical and oil-diffusion pumps. Typical run

times ranged between 30 and 120 milliseconds, which required ultra-high-speed

instrumentation and control equipment. Data was recorded using high-speed oscillographs

and magnetic tape.

This tunnel provided Langley researchers with experience operating non-

conventional tunnels, and modifications eventually allowed some tests above Mach 30.

While the results were useful, the high energy involved with the arc heater continued to pose

problems and risk. After an accident seriously damaged this tunnel, it was demolished in

1969. No portion of it is known to have survived.14

15-Inch Low-Turbulence Tunnel

The 15-inch Low-Turbulence Tunnel is a self-contained, closed-circuit tunnel that

can produce very stable air flows up to about 88 miles per hour. Its test section consists

primarily of clear panels, and it offers excellent optical access for laser velocimetry (LV)

and particle-image velocimetry (PIV) measurements. Recent research involving active flow

control devices to improve jet engine inlet performance in adverse conditions has relied

heavily of the tunnel’s PIV capabilities. The tunnel remains in service.15

D-104 – 20- by 28-Inch Shear-Flow Tunnel

The 20- by 28-Inch Shear-Flow Tunnel is an atmospheric, open-circuit tunnel with a

top speed of approximately 100 miles per hour that was purchased as a self-contained unit

from Aerolab about 19__. Intended primarily for fundamental boundary-layer research, a

series of inlet screens and a large contraction ratio deliver air to the test section with very

low freestream turbulence. The 15-foot-long test section features large windows on both

sides and an adjustable ceiling for control of the pressure gradient through the full length of

the test section. The large windows provide excellent optical access for laser velocimetry

(LV) and particle-image velocimetry (PIV) measurements. This tunnel remains in service.16

D-105 – 7- by 11-Inch Low-Speed Tunnel

The 7- by 11-Inch Low-Speed Tunnel is an atmospheric, open-circuit tunnel with a

top speed of approximately 100 miles per hour that was purchased as a self-contained unit

from Aerolab about 1965, initially installed in room B-115, and moved into room D-105

sometime later. (Room D-105 earlier housed a 20-Megawatt Linear Plasma Accelerator

used for magnetoplasmadynamics studies.) Initially intended as a pilot for the 14- by 22-

Foot Subsonic Tunnel, it has been primarily utilized for fundamental research since 1970. It

is now optimized for direct-drag measurements of flat panels, and it employs a linear air

bearing drag balance. This tunnel is equipped with a variety of pitot and hotwire-type flow

measurement instruments, as well as smoke-injection and smoke-wire capabilities for flow

visualization. Its primary use has been for studies into drag reduction within turbulent

boundary layers, but it achieved some unusual notoriety in 2008 when swimmers Michael

Phelps and Natalie Coughlin made Olympic history wearing swimsuits made of a drag-

reducing fabric developed in part through testing in this tunnel by Langley researcher

Stephen Wilkinson. The tunnel remains in service.17

D-106 – 12-Inch Mach 6 High-Reynolds Number Tunnel

Also known as the Mach 6 High-Pressure Tunnel, the 12-Inch Mach 6 High-

Reynolds Number Tunnel entered service in 1967. It was a conventional tunnel that used

heated dry air furnished by the facility and exhausted through an aftercooler into a 41-foot-

diameter vacuum sphere to generate a speed of Mach 6. Electric resistance-type heaters

heated the incoming air to a maximum temperature of about 540 Fahrenheit. This tunnel

featured an axisymmetric nozzle into two interchangeable, 12-inch-diameter test sections.

One was equipped with schlieren windows and a model-injection mechanism capable of

handling models up to 4-feet long. The second test section was designed for investigations

of boundary layers along the tunnel wall using a variety of measurement techniques. The

tunnel was closed about 1998 and demolished in 2007 to make room for two acoustic

research devices being relocated from other buildings in the West Area, the Curved-Duct

Test Rig from Building 1218, and the Grazing-Flow Impedance Tube from Building 1287.18

D-107 – 20-Inch Supersonic Wind Tunnel

The 20-Inch Supersonic Wind Tunnel (SWT) was originally designed and built at the

Jet Propulsion Laboratory (JPL) in Pasadena, California, in 1948, a full decade before JPL

became part of NASA. At the time, JPL, a unit of the California Institute of Technology,

was heavily engaged in rocket engine development, and this tunnel was one of two it built to

support that activity. After JPL became a unit of NASA, its hardware role gradually shifted

from rocket engines to the development of unmanned spacecraft, drastically reducing its

need for wind tunnels. By 1980, the laboratory needed the space they occupied for other

facilities, so NASA looked to relocate them within the agency.

Robert L. Trimpi, a Langley researcher, realized that the 20-inch JPL tunnel could be

useful at Langley, and he recommended that it be moved. Actually, only the core of the

tunnel—its nozzle, test section, and diffuser—were moved to Virginia. JPL had powered

the tunnels with dedicated air compressors that were not needed at the HFC, which could

also furnish vacuum at the exhaust end. Between 1982 and 1986, Trimpi oversaw its move

and installation in room D-107. A new air-inlet section, enclosed in a room that filled part

of the central hall, was built, and the diffuser exhaust was mated to a new duct with a valve

that directed the flow either to atmosphere or to a 60-foot-diameter vacuum sphere. Langley

named it the 20-Inch Supersonic Wind Tunnel.

This tunnel had a unique distinction. It was the first supersonic tunnel built that used

flexible top and bottom walls to vary the speed. Supersonic nozzles must be precisely

contoured for a specific speed, but JPL engineers had devised a nozzle that used a series of

jacks to flex sheets of stainless steel and, thus, re-contour the nozzle for a range of speeds

between Mach 1.4 and 5.0. Both sides of the nozzle retract for access to the nozzle and the

forward portion of the 20- by 18-inch test section. Trimpi added an 8-foot-diameter settling

chamber, along with a “quiet” valve and flow-conditioning section to supply stable air to the

nozzle and, thus, help create a uniform flow through the core of the test section.

Models are typically sting-mounted on a support that allows pitch adjustment from

34 to -9.5 degrees and ±8.5 degrees in yaw. They are protected from startup and shutdown

transient temperatures and loads by a plenum isolation door and a rapid injection capability

of the support system. A variety of heat-transfer and static pressure instrumentation is

available, plus laser velocimetry for flow-field measurements. In recent years researchers

developed a way for the tunnel to operate subsonically as well, and it has proven useful for

airfoil testing in the Mach 0.35 – 0.75 range. Maximum run times approach 30 minutes

when the exhaust is to atmosphere. The tunnel received several modifications in 1994 to

improve its reliability and ease of use, including a connection to the HFC’s 100-foot-

diameter vacuum sphere that enables it to use the total volume of the 60- and 100-foot

spheres for longer run times. The 20-Inch Supersonic Wind Tunnel has been a very

productive facility at Langley, and it remains in active service.19

D-108 – 20-Inch Mach 6 Tunnel

The 20-Inch Mach 6 Tunnel, opened in 1958 as the 20-Inch Hypersonic Tunnel, is a

conventional tunnel utilized for aerodynamic and aerothermodynamic tests of proposed

aerospace vehicles, and for the exploration of basic fluid dynamic phenomena, including

boundary layer laminar-to-turbulent transition. The tunnel uses dry air furnished by the

facility and heated by an electric resistance heated, and it exhausts through an aftercooler

into a vacuum sphere, to the atmosphere (assisted by an air ejector), or a combination of the

two. Its diffuser includes a moveable second minimum. The tunnel delivers a speed of

Mach 6, with stable flow in a core test area measuring 12-inches square in its 20- by 20-inch

test section. Models can be fixed on the floor, or sting mounted on an injection mechanism,

which allows a pitch setting from 55 degrees to -5 degrees and a yaw range of 0 to -10

degrees. Run times vary from 2 to 20 minutes, depending on the choice of exhaust.

The results of tests conducted in this tunnel are often compared to those from tests in

the 31-Inch Mach 10 and 20-Inch Mach 6 CF4 tunnels to more fully assess the effects of

compressibility and real-gas aerothermodynamics under similar Reynolds number

conditions. This tunnel performed a large number of tests during the Space Shuttle

development program, and it is kept available during missions to investigate potential

problems. The 20-Inch Mach 6 Tunnel is in service.20

D-114 –Mach 6 Pilot Quiet Tunnel

Also known as the Nozzle Test Chamber, this facility was not actually a research

tunnel per se, but rather a conventional device used primarily to develop a low-disturbance,

or quiet, hypersonic nozzle. Constructed about 1995, and originally installed in room D-

108, it received heated air from the building’s central system and exhausted through an

aftercooler to a vacuum sphere. The extended settling chamber ahead of the asymmetric

nozzle measured 85 inches long and contained as many as 11 screens, including one formed

out of steel wool, to eliminate turbulence. The nozzle contraction also featured an annular

slot connected to a vacuum source to remove any remaining boundary-layer immediately

ahead of the throat. The 19-inch-diameter nozzle exhausted into a larger test section, where

an overhead model-injection device inserted the model once stable flow had been achieved.

Air exited the test section through a conical collector and diffuser to a vacuum vessel.

While complex to build and operate, the apparatus achieved very-smooth, low-turbulence

flow at Mach 6, and its success has prompted additional developmental work by others in

the field. This tunnel was relocated to room D-114 in the late-1990s, after the removal of

other equipment freed the space. After almost a decade of service, the Mach 6 Pilot Quiet

Tunnel was dismantled in 2005 and donated to Texas A&M University in College Station,

Texas, where it is currently being modified for a return to service.21

D-115 – 18-Inch Mach 8 Variable-Density Tunnel

Opened in 1960, the 18-Inch Mach 8 Variable-Density Tunnel, also known as the

Mach 8 Variable-Density Hypersonic Tunnel and the 18-Inch Mach 8 Tunnel, is a

conventional tunnel capable of producing a wide range of Reynolds numbers. It uses dry air

heated by a two-stage heater that a control valve delivers to the settling chamber at 20 to

3,000 pounds per square inch. The heater’s first stage employs a circulating ethylene glycol

fluid that heats the air to approximately 150 Fahrenheit. An electric resistance heater then

raises the temperature to as high as 1,030 Fahrenheit. The settling chamber contains screens

to stabilize the flow, which then passes through the axisymmetric nozzle and 18-inch-

diameter test section. Depending on the operating pressure, the core test area measures

between 4 and 14 inches in diameter. A model injection system raises the model, which can

be up to 27 inches long, into the stabilized air stream. The sting has adjustments for pitch

and yaw. The tunnel exhausts through a diffuser with a second minimum to the atmosphere

or into a pair of vacuum spheres, 41 and 60 feet in diameter. This tunnel produces speeds

between Mach 7.5 and 8.0, depending on pressure. The maximum run time to vacuum is

about 90 seconds, but runs as long as 10 minutes are possible when exhausting to the

atmosphere. This tunnel has largely been on standby since the late-1970s, but it remains

intact.22

D-116 – 20-Inch Mach 8.5 Tunnel

Opened in 1961, the 20-Inch Mach 8.5 Tunnel was a conventional tunnel that

generated the highest speed using air of all Langley tunnels. It used dry air heated by an

electric resistance heater to 1,000 Fahrenheit. The settling chamber contained screens to

stabilize the flow, which then passed through the axisymmetric nozzle and 21-inch-diameter

test section. (Why “20-Inch” was chosen for its name is unknown.) The core test area

measured 16 inches in diameter. A model injection system raised the model into the

stabilized air stream. The sting had adjustments for pitch and yaw. The tunnel exhausted

through a diffuser with a movable second minimum to the atmosphere or a vacuum sphere.

The maximum run time to atmosphere was about 7 minutes and about 30 seconds to

vacuum. This tunnel was demolished about 1980 to provide space for the Pilot Supersonic

Low Disturbance Tunnel.23

6- by 10-Inch Supersonic Low-Disturbance Tunnel

Originally built in 1988 as the Pilot Supersonic Low Disturbance Tunnel, this facility

is a Mach 3.5 conventional tunnel with specific features intended to produce a low-

turbulence, or “quiet” flow through the core of its test section. Using air as the test gas, its

quiet flow is achieved using a combination of settling chamber devices, a boundary-layer

extraction system, and a highly-polished, 2-dimensional nozzle with a boundary-layer bleed

upstream of its throat. Although a boundary does form in the nozzle that ultimately

generates a shock wave and downstream turbulence, the design produces a very quiet core in

the shape of a rhombus 5 – 10 inches by 1.5 – 3.6 inches in size (the size varies with

Reynolds number). Since this tunnel exhausts to the atmosphere, typical run times range

from 15 to 30 minutes. This tunnel’s primary use has been in developing supersonic quiet-

flow technology and computational methods. It currently sees limited use, but remains

intact.24

D-118 – Probe-Calibration Tunnel

The Probe Calibration Tunnel is a conventional, open-jet pressure tunnel with three

interchangeable nozzles, and it has adjustments for the independent control of pressure,

temperature, and speed. Test section pressure can be varied from 0.13 to 10 atmospheres (2

– 146 psia), and stagnation temperature can be set from 0 to 200 Fahrenheit. Its three

interchangeable nozzles provide a speed range of Mach 0.05 – 3.5. As its name indicates,

this tunnel is used primarily to calibrate instrumentation used in the Hypersonic Facilities

Complex’s other tunnels. This tunnel is in service.25

Another related tunnel is considered part of the Hypersonic Facilities Complex

although it is located in another West Area building some distance from Building 1247:

Building 1275 – 20-Inch Mach 6 CF4 Tunnel

Previously known as the Hypersonic CF4 Tunnel, this conventional tunnel was the

result of an extensive rebuilding of the 20-Inch Hypersonic Arc-Heated Tunnel in the mid-

1970s. Where the original tunnel tested materials at high temperatures, the rebuilt facility

was designed to investigate high-energy aerothermodynamic phenomena, particularly those

around a blunt body during atmospheric entry. The extreme temperature and shock induce

molecular dissociation. When this occurs, the specific properties of the “real” gas deviate

significantly from those of a so-called “ideal” gas. While impulse-type tunnels can generate

these high-enthalpy conditions momentarily, their data suffer from poor test repeatability.

Helium allows high-density shocks to be created in conventional tunnels at relatively low

enthalpy, but its real-gas properties differ significantly from those of air.

The researchers looked to tetrafluoromethane (CF4) as a test gas because it

dissociates in a manner very similar to air, but at lower enthalpies, allowing a conventional

tunnel to reproduce the real-gas phenomena associated with hypersonic flight at reasonable

temperatures. The 20-Inch Mach 6 CF4 Tunnel was the result, and it is currently the only

operational tunnel in the United States with this capability.

The CF4 Tunnel is similar to the 20-Inch Mach 6 Tunnel for air, and its performance

envelope complements it, making possible comparative analyses of data from the two

tunnels to establish the magnitude of real-gas effects at constant Mach and Reynolds

numbers. The CF4, at pressures up to about 2,500 pounds per square inch, is heated by two

lead-bath heaters to approximately 1100 Fahrenheit and passes through screens in a settling

chamber, through the axisymmetric Mach 6 nozzle, and into the test section. The test

section is an 8- by 5-foot box through which the open jet flows. From there, the gas passes

through a diffuser and aftercooler to a pair of vacuum spheres. It is equipped for schlieren

photography; force, moment, and pressure measurements; and heat transfer studies. Runs

typically last 30 seconds or less, after which on-site equipment recovers the CF4 for re-use.

The CF4 Tunnel has undergone several modifications over the years to improve its

performance, and it remains in service.26

Notes 1 Donald D. Baals and William R. Corliss, Wind Tunnels of NASA (Washington: National Aeronautics and Space Administration, 1981), 55-56. For a more rigorous discussion of hypersonic aerodynamics, see John D. Anderson, Hypersonic and High Temperature Gas Dynamics (New York: McGraw-Hill, 1989, reprinted by Reston, VA: The American Institute of Aeronautics and Astronautics, 2000), 13-23. 2 James R. Hansen, Engineer in Charge: A History of the Langley Aeronautical Laboratory, 1917-1958 (Washington: National Aeronautics and Space Administration, 1987), 347. Also Baals and Corliss, Wind Tunnels of NASA, 59-60. 3 Baals and Corliss, Wind Tunnels of NASA, 60. 4 Hansen, Engineer in Charge, 349-350. 5Ibid, 350-353. 6 In addition to sources for individual tunnels cited, information for these tunnels was compiled from Langley Real Property Office records for Building 1247 (A – J). These include “Real Property Record-Buildings,” Form 845; “Real Property Record-Other Structures and Facilities,” Form 846; “Real Property Transaction Vouchers,” Form 1045; “General Ledger,” GAO Standard Form 1014; and other untitled papers in the folders for this building. Among these papers was W. Ray Hook and Roy V. Harris, Jr. to Head, Real Estate Management Office, ”Name Change for Wind Tunnels in Hypersonic Facilities Complex,” Oct. 5, 1992. 7 William T. Schaefer, Jr., “Characteristics of Major Active Wind Tunnels at the Langley Research Center,” Technical Memorandum X-1130 (Washington: National Aeronautics and Space Administration, 1965), 23. Also Martin A. Weiner, “Resume of Research Facilities at the Langley Research Center,” (Hampton, VA:

NASA Langley Research Center, July 1968), pages not numbered. Also John Warren, HFC Control Room Technician, conversation with author, Feb. 27, 2009. 8 C. G. Miller and F. M. Smith, “Langley Hypersonic Facilities Complex – Description and Application,” Paper No. 86-0741-CP (Reston, VA: The American Institute of Aeronautics and Astronautics, 1986), 5-6. Also John Warren, HFC Control Room Technician, conversation with author, Feb. 27, 2009. 9 Ralph D. Watson and Dennis M. Bushnell, “Calibration of the Langley Mach 20 High Reynolds Number Helium Tunnel Including Diffuser Measurements,” Technical Memorandum X-2353 (Washington: National Aeronautics and Space Administration, 1971), 3-4. 10 Weiner, “Resume of Research Facilities at the Langley Research Center,” pages not numbered. Also R. D. Watson, D. J. Morris, and M. C. Fischer, “Operating Characteristics of the Langley Mach 10 High Reynolds Number Helium Tunnel,” Technical Memorandum X-2955 (Washington: National Aeronautics and Space Administration, 1974), 3-4. 11 “AAAC Research Facilities,” AAAC Annual Report, FY 2002-2003, (Hampton, VA: NASA Langley Research Center, Aerodynamics, Aerothermodynamics and Acoustics Competency, 2003), 122. Also “Expansion of Basic Aerodynamics Research Facility,” CoF Project Initiation Form originated by P. M. Jackson, Feb. 25, 1997, Langley Real Property Office records for Building 1214. 12 Weiner, “Resume of Research Facilities at the Langley Research Center,” pages not numbered. Also Ivan E. Beckwith, William D. Harvey, and Frank L. Clatk, “Comparisons of Turbulent-Boundary-Layer Measurements at Mach Number 19.5 with Theory and an Assessment of Probe Errors,” Technical Note D-6192 (Washington: National Aeronautics and Space Administration, 1971), 8. 13 “AAAC Research Facilities,” 143. Also William B. Boatright, Alexander P. Sabol, Daniel I. Sebacher, Shimer Z. Pinckney, and Robert W. Guy, “Langley Facility for tests at Mach 7 of Subscale, Hydrogen-Burning, Airframe-Integratable, Scramjet Models,” Journal of Aircraft, vol. 13, no. 2, Feb. 1976, 67. Also R. Wayne Guy, R. Clayton Rogers, Richard L. Puster, Kenneth E. Rock, and Glenn S. Diskin, “The NASA Langley Scramjet Test Complex,” Paper No. 96-3243 (Reston, VA: The American Institute of Aeronautics and Astronautics, 1996), 5-6. 14 Weiner, “Resume of Research Facilities at the Langley Research Center,” pages not numbered. Also Fred M. Smith, Edwin Harrison, and Pierce L. Lawing, “Description and Initial Calibration of the Langley Hotshot Tunnel with Some Real-Gas Charts for Nitrogen,” Technical Note D-2023 (Washington: National Aeronautics and Space Administration, 163), 3-8. Also Charles G. Miller III, “Langley Hotshot Tunnel Operations with Helium at Mach Numbers in Excess of 30,” Technical Note D-5901 (Washington: National Aeronautics and Space Administration, 1970), 5-6. 15 “AAAC Research Facilities,” 120. Also Susan A. Gorton, “Active Flow Control for Inlets,” Agenda and Abstracts, NASA Glenn Research Center UEET (Ultra-Efficient Engine Technology) Program, 35. 16 “AAAC Research Facilities,” 121. 17 “AAAC Research Facilities,” 125. Also Patricia Phillips, “Michael Phelps, Sports illustrated, & NASA Technology: The Geek & the Gold,” www.examiner.com/x-504-Space-News-Examiner~y2008m8d19-Michael-Phelps-Sports-Illustratedand-NASA-Technology, Aug. 19, 2008. 18 Weiner, “Resume of Research Facilities at the Langley Research Center,” pages not numbered. Also John Warren, HFC Control Room Technician, conversation with author, Feb. 27, 2009. 19 “AAAC Research Facilities,” 123. Also J. L. Dillon, R. L. Trimpi, and A. E. Schwartz, “The Langley 20-Inch Supersonic Wind Tunnel,” Paper No. 86-0765-CP (Reston, VA: The American Institute of Aeronautics and Astronautics, 1986), 4-6. Also Pete Jacobs and Lori Rowland, “20-Inch Supersonic Wind Tunnel: NASA

Langley Research Center,” http://wte.larc.nasa.gov/facilities_updated/fluid_dynamics/20inch_supersonic2.htm, Oct. 25, 2006. 20 “AAAC Research Facilities,” 140. Also Theodore J. Goldberg and Jerry N. Hefner, “Starting Phenomena for Hypersonic Inlets with Thick Turbulent Boundary Layers at Mach 6,” Technical Note D-6280 (Washington: National Aeronautics and Space Administration, 1971), 14-15. Also John Warren, HFC Control Room Technician, conversation with author, Feb. 27, 2009. 21 “NASA Langley Mach 6 Quiet Tunnel,” Texas A&M University Department of Aerospace Engineering, flight.tamu.edu/tunnel/mach6quiet.html, Feb. 4, 2009. Also John Warren, HFC Control Room Technician, conversation with author, Feb. 27, 2009. 22 C. G. Miller and F. M. Smith, “Langley Hypersonic Facilities Complex – Description and Application,” Paper No. 86-0741-CP (Reston, VA: The American Institute of Aeronautics and Astronautics, 1986), 4. Also William T. Schaefer, Jr., “Characteristics of Major Active Wind Tunnels at the Langley Research Center,” Technical Memorandum X-1130 (Washington: National Aeronautics and Space Administration, 1965), 17. Also John Warren, HFC Control Room Technician, conversation with author, Feb. 27, 2009. 23 Weiner, “Resume of Research Facilities at the Langley Research Center,” pages not numbered. Also William T. Schaefer, Jr., “Characteristics of Major Active Wind Tunnels at the Langley Research Center,” Technical Memorandum X-1130 (Washington: National Aeronautics and Space Administration, 1965), 20. Also John Warren, HFC Control Room Technician, conversation with author, Feb. 27, 2009. 24 “AAAC Research Facilities,” 124. Also Robert E. Bower, “Current Wind Tunnel Capability and Planned Improvements at the NASA Langley Research Center,” Paper No. 86-0727-CP (Reston, VA: The American Institute of Aeronautics and Astronautics, 1986), 52-53. 25 “AAAC Research Facilities,” 126. Also John Warren, HFC Control Room Technician, conversation with author, Feb. 27, 2009. 26 “AAAC Research Facilities,” 141. Also John R. Micol, Raymond E. Midden, and Charles G. Miller III, “Langley 20-Inch Hypersonic CF4 Tunnel: A Facility for Simulating Real-Gas Effects,” Paper No. 92-3939 (Reston, VA: The American Institute of Aeronautics and Astronautics, 1992), 2-4. Also Raymond E. Midden, and Charles G. Miller III, “ Description and Calibration of the Langley Hypersonic CF4 Tunnel: A Facility for Simulating Low γ Flow as Occurs for a Real Gas,” Technical Paper 2384 (Washington: National Aeronautics and Space Administration, 1985), 1-4.