THE HIGH SPEED FRONTIER
 
 
Chapter 3: Transonic Wind Tunnel Development (1940 -1950)
 
THE REPOWERED 8-FOOT HIGH-SPEED TUNNEL; SMALL MODEL TECHNIQUES
 
 
 
[68] The original 8000-hp drive of the 8-foot tunnel produced maximum test speeds with typical models of about Mach 0.75, and the bulk of the testing done in this tunnel was limited in speed not by choking but by lack of power. When Stack became section head in 1939 about three years after the 8-foot began operating, he almost immediately started talking about the need to increase the power. He had been accustomed to no power limitations in the 11-inch and 24-inch tunnels and tests of airfoil models in those facilities usually extended upward to the choking limit. At first we thought in terms of enough power to provide Mach 1 in the empty 8-foot tunnel plus a margin for installation of models, giving a total requirement of 12 000 hp. As time went on, however, the need for testing at low supersonic speeds became more apparent, and by 1942 when the first tentative Langley management approvals of a repowering plan were obtained, 18 000 hp had been decided on (ref. 54), a rather arbitrary increase of 50 percent over the original figure. (A later agency press release has it that this liberal level of power was in anticipation of the large requirements for a ventilated type transonic tunnel. Actually, it was based on the idea of achieving a supersonic operating capability for which the power requirements [69] were uncertain at that time.) Little is to be found in the way of documentation relating to the promotions of the repowering up to 1942. For the most part, it was talk between Stack and his bosses, Miller and Reid. Occasionally, Stack dashed off handwritten notes to Miller which did not survive in the Langley files. There were no exhaustive reviews by any advisory groups as there would have to be today. Actually there was practically no substantive concept development or design study behind all the talk up to 1942; we never made any engineering designs of model support systems or test section modifications for supersonic testing during this period. It was simply taken for granted that all of this would be done later if plans for an ample power increase went through. The proposal went all the way to Lewis and finally gained his approval, supported mainly by informal discussions and general good intentions to work on the problems later.
 
The 16-foot high-speed tunnel which started operating in 1941 had been built with a 16 000-hp drive, the maximum power available at Langley at that time. More seriously underpowered than the 8-foot tunnel, it could reach a maximum speed of only about Mach 0.7 with the smallest model test setups. The principal use of this tunnel as originally conceived was to extend the kind of full-scale propeller and engine nacelle-propeller testing done in the old Propeller Research Tunnel to high speeds. After the tunnel was well along in construction, it became clear that full-scale engine nacelles would produce such enormous blockage effects that choking would occur typically at speeds as low as Mach 0.6, and that throughout the entire speed range major distortions would be present in the data. Only a few such setups were tested during the first years of operation, primarily to investigate and improve radial-engine cooling.
 
Early in 1943 Reid and Miller decided to create a new research division to incorporate all the ground-based high-speed aerodynamics activities including the following groups: 8-foot tunnel, 16-foot tunnel, 9-inch supersonic tunnel, and the group under A. Kantrowitz involved with fundamental as dynamics research.
 
The new division was called "Compressibility Research," compressibility being a basic property of gases which becomes important in aerodynamics at high speeds. (Langley usually favored vague general [70] organizational titles, believing they might help to discourage criticism of what was going on, and insure that researchers were not unduly hemmed in by nominal organizational boundaries.)
 
Stack's first problem as a new division chief in mid-1943 was to appoint a replacement for David Biermann who was vacating his position as head of the 16-foot tunnel section for industrial employment. He selected me for the job, but I was not happy about it. I was comfortable with the 8-foot group, and we were in the midst of promising plans for the repowering. By comparison, it seemed to me that the underpowered 16-foot was doomed to routine testing at subcritical speeds. Stack's answer to these misgivings was, "By God, we'll repower 16-foot!" He was elated at this first contemplation of an exciting new crusade, and I was sufficiently encouraged to move into my new assignment with some enthusiasm in July 1943. One of the first visitors to my new office was Mr. Miller. He emphasized the importance of the job and offered some typical advice, "Don't do anything without first checking with Stack or me."
 
In my last weeks at 8-foot, I had started work on the problem of how best to support test models in the repowered tunnel to provide testing as close as possible to Mach 1. Byrne's results (ref. 87) gave a firm indication of how small the test models would have to be, and it was obvious from the outset that conventional strut supports (fig. 13) could not be used because the struts themselves would contribute more blockage than the small test models. Upon moving over to 16-foot, I continued to study this problem as time permitted, partly because I knew that we would eventually be confronted with it when the 16-foot was repowered, but mainly because I had developed an interest in it.
 
For wing testing, I first considered half-span models mounted from the tunnel wall. This eliminated the struts, but the tunnel wall boundary layer, several inches thick, made the flow over the root section of the wing invalid-an especially serious deficiency for the small wings that would have to be used. Mounting the wings on a support plate which bypassed the wall boundary layer was considered next, but the asymmetry of this arrangement seemed clearly to be undesirable at near-choking speeds. And then a symmetrical solution suggested itself: locate the support plate in the center of the tunnel in the plane of....
 

photo of strut supports inside wind tunnel
 
[71] FIGURE 13.-Typical strut support system used prior to 1944 in the 16-Foot High-Speed Tunnel. Choking speed was about Mach 0.8.
 
....symmetry of a complete wing model (fig. 14, top). The thickness of the plate would have no effect on choking because in effect a new strutless test section was established on each side of the plate. Being at the plane of symmetry of the wing, the plate would not affect the wing flow, and the plate boundary layer was negligibly thin.
 
I recall a sense of satisfaction as I described the center-plate support idea to Stack in mid-1944 during one of his frequent visits to my office in the 16-foot tunnel building. He proposed to start design work at once to implement the idea in the 8-foot tunnel, which was to shut down for repowering in a few months. Care was taken by the design group to shape the leading edge of the plate so as to avoid a local velocity peak. By the time 8-foot commenced operations with its new 18 000-hp drive in February 1945, the center plate was ready for installation. The first wings tested were part of a comprehensive general research program set up in November 1944 to support the Army Air Corps' first jet-powered high-speed bomber development. Wings of...
 

cross-sectional view of blade shapped support inside wind tunnel
 
[72] FIGURE 14.-Support systems developed by Langley which do not cause a decrease in choking Mach number. From a 1946 Conference chart.
 

photo of center support blade  installed in wind tunnel
 
[73] FIGURE 15.-Center-plate support for wing testing installed in the repowered 8-Foot High-Speed Tunnel, 1945.

 
[74] 6-inch root chord, 38-inch span, and 10 percent thickness ratios made the choking speed about Mach 0.95; reliable data were obtained up to about Mach 0.93 (see fig. 15). The first comprehensive pressure distributions and wake-survey drag measurements were transmitted to the Army in the summer of 1945 (refs. 88, 89) and other wing and tail configurations followed in short order (ref. 90).
 
I have described the center-plate development in some detail because it is a good example of innumerable creditable but unspectacular contributions made by NACA supervisors and seasoned researchers, routinely and usually anonymously. So much of this happened all the time that it would have been quite impractical for NACA authors to acknowledge all such contributed ideas in their papers. If an idea had been formalized by a memorandum to the Chief of Research, by a patent application, or by a publication, some acknowledgement would be expected; but in the absence of such documentation, the origin was likely to be quickly forgotten. In the case of the center plate, the test reports give only a description of the device, and Stack's later brief review of facilities (ref. 54) says only that it evolved from "intensive study."
 
As we progressed in transonic research, we learned that the prime problems lay not with isolated wings but with wing/bodies and complete configurations. The center plate was not well adapted for testing such configurations, and some type of sting support system was needed. In this case, the support-choking problem was not the sting itself but the large strut downstream of the test model which extended to the tunnel walls or to an external balance. The avoidance of strut choking in this setup had a more obvious solution: divergence of the tunnel walls to compensate for the strut area blockage. If the strut were located in the test section, this would have required a major mechanical operation on the tunnel structure, and it was easy to see that the same effect could be realized much more expediently simply by installing an insert or liner within the existing walls to create a new throat section for the test model ahead of the sting support strut (fig. 14, center). The same scheme could obviously also be used to avoid strut choking for the propeller dynamometer installations in the 8-foot and 16-foot tunnels (fig. 14, bottom). The principle of these liners seems to have evolved from informal group discussions in 1944. E. C. Draley, R. H. Wright, [75] E. Palazzo, and R. Moberg were among the implementers of the new support systems. Initially, the sting was attached to a balance outside the tunnel through a large strut housed within a fairing. This produced troublesome tare forces, and a much improved arrangement used small strain-gage balances contained within the test models. With this latter arrangement the support strut could be located farther downstream in the diverging diffuser section where it would not contribute to choking in the test section. Thus the effect of the liner (fig. 14, middle sketch) was achieved without any alteration to the tunnel contours.
 
The D-558-1 and the X-1 research airplanes were the first configurations tested with the new sting systems (fig. 16), providing extremely important data at speeds up to about Mach 0.92 prior to the first high-speed flights of these aircraft.
 
In December 1947, the 8-foot tunnel test section was equipped with a plaster throat insert contoured theoretically to produce uniform shockless flow at Mach 1.2 (ref. 93). The nozzle shape was perfected experimentally by tracing pressure disturbances measured near the tunnel center line back to their point of origin on the wall, and then...
 

photo of X-1 model supported in wind tunnel
 
FIGURE 16.-Sting-supported model of the X-1 in the repowered 8-Foot High-Speed Tunnel, 1946.

 
[76] making the minute changes in wall contour determined theoretically to be needed (ref. 54). Once perfected, the nozzle was used for the remaining two years of closed-throat operation, finally being removed in January 1950 to make way for the slotted test section.
 
It was concluded from the success of the Mach 1.2 closed-throat nozzle that the same techniques could be applied at lower supersonic speeds; Mach 1.1 was thought to be clearly feasible. The problem of reflection of the bow shock back on to the rear of the test model would probably determine the lower limit rather than any limitation of nozzle design.
 
Thus by seriously coming to grips with the choking problem, NACA work in the early forties reduced the unattainable speed range for closed tunnels to the narrow region between Mach 0.95 and about 1.1, approximately one-third its former proportions. The price that had to be paid for the small-model technique was, of course, a reduction in test Reynolds number. Even so, test Reynolds numbers of the order of one-fifth those of the small research airplanes could be obtained, close enough to permit very important valid comparisons.
 
One of our most important duties as NACA supervisors was to insure the prompt flow of the results of our research to industry and the military. NACA had learned by hard experience in the twenties that the issuance of technical documents, while of course essential, was not sufficient as the sole mode of communication. The top managers in industry and in the military seldom had time to read NACA technical reports, and-equally important from NACA's viewpoint-Congressmen had neither the time nor the qualifications to read the technical reports and judge whether the agency's output justified its appropriations. Starting in 1926, the so-called Engineering Conferences provided periodic opportunities to highlight recent research accomplishments, and at the same time to "blow the horn" for the agency in a most effective and unobnoxious way. Great care was taken to make these presentations simple enough for managers and Congressmen to understand without losing any important technical implications.
 
In 1946 it was especially important to reveal and advertise our progress in transonic and supersonic testing capabilities. We spent some time developing conventional charts and illustrations, but I was unhappy [77] with the rather uninspiring results. Finally, we decided to replace the charts and pictures with live action. We built two small wind tunnels with 6-inch glass-sided throat sections revealing not only the tunnel contours but also schlieren images of the shocks formed on the test models and in the tunnel diffusers (fig. 17). Above the tunnels was a manometer board calibrated to show the velocities along the tunnel walls and over the test models. The lower tunnel was a conventional...
 

cross-sectional view of a small wind tunnel
 
FIGURE 17.-Small tunnels used to demonstrate choking, supersonic nozzles, power requirements, and transonic and supersonic airfoil flows, 1946.

 
....subsonic design which illustrated choking and the upper was a convergent-divergent supersonic tunnel illustrating the principles of NACA's three large new supersonic facilities then in design. Because of our very limited budget, we had to employ a unique drive system for the little tunnels: they were connected by a long diffuser to the low-pressure test chamber of the 16-foot high-speed tunnel. By running the 16-foot at something over 400 mph, enough suction was provided to choke the subsonic model tunnel and generate Mach 1.6 in the supersonic tunnel. Most of the visitors, hearing only a distant rumble, were not aware that the 16-foot was being used as an oversized pump to activate the little tunnels.
 
The first demonstrations were made in January 1946 for a conference of aviation writers, and their reactions provided ample evidence of the effectiveness of this show. After the little tunnels had served again in the 1946 Annual Inspection of the Langley Laboratory, we used the supersonic tunnel to investigate the flow phenomena and forces on a control flap at supersonic speeds (ref. 95), the first time this problem had been examined in a supersonic tunnel. Schlieren photographs taken in these demonstration tunnels also found their way into several books and periodicals.
 

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