-
THE HIGH SPEED
FRONTIER
-
-
- Chapter 2: The High-Speed
Airfoil Program
-
- SUPERCRITICAL AND TRANSONIC
AERODYNAMICS (1945-1956)
-
-
-
- [36] The emergence of
the transonic research airplanes in the mid-forties (Chapter III)
greatly heightened our interest in the supercritical behavior of
airfoils and in developing testing techniques for exploring the
supercritical and transonic regions. But an even stronger
motivation had developed concurrently with the advent of the
turbojet engine. In 1944, the Army had sent an XP-59A, the first
U.S. jet-powered airplane, for flight demonstrations at Langley.
Standing beside the main runway and watching this airplane fly by
at nearly 400 mph, we sensed for the first time that here was the
key to transonic and supersonic flight, a practical new propulsion
concept capable of the enormous power required to penetrate the
transonic region. The so-called "sound barrier," which
had been almost universally thought of as a set of
adverse aerodynamic problems, in reality also involved a
fundamental limitation of the piston engine due to its fixed
"displacement," or capacity to inhale air for combustion. Since
the displacement was independent of airspeed, no significant
increase in peak power could occur as the flight speed increased.
Thus, there had been no realistic hope that piston engines could
be developed in the sizes that would have been needed for
transonic flight; the transonic "barrier" was actually as much a
piston engine barrier as an aerodynamic barrier. The jet engine on
the other hand ingests a volume flow of air that increases as
the flight speed [37] increases, permitting a continuous increase in
power in contrast to the fixed power of the piston engine. This
power increase is significantly augmented at high speeds by the
"ram" pressures of the air which provides supercharging and
improves the cycle efficiency.
-
- We understood the principles and enormous
potential of the turbojet only vaguely at the time of the XP-59A
demonstration. Very little data were available to us on the
details and performance of the G.E. I-16 engine. Several of us
spent the next few days in exciting speculations of possible
jet-engine thermodynamic cycles, airflow characteristics, and
crude performance estimates, which gave us a better understanding.
K. F. Rubert, who had taught internal combustion engines at
Cornell, undertook a more careful systematic analysis, published
in 1945 in a paper which I reviewed as chairman of the Langley
Editorial Committee (ref. 44). (Periodic editorial duties of this kind were of
great value as a means
of education and stimulation of all
involved-in addition to their obvious direct benefit to the
quality and accuracy of Langley reports.)
-
- By now our limited goal of the 550-mph
subcritical airplane of the mid-thirties had become meaningless
and we could foresee the imminent achievement of supersonic
flight. Few doubted that operational supersonic military aircraft
would soon follow the research airplanes. The need to acquire
accurate supercritical and transonic aerodynamic data had become
acute, and Langley researchers responded to the challenge with
considerable inventiveness. Eight innovative techniques were
eventually devised and explored in various forms by NACA, ending
with general acceptance of the semi-open tunnel for
two-dimensioned airfoil testing up to Mach 1, and the slotted
transonic tunnel for wing and aircraft configuration testing as
the most satisfactory devices. (These developments are listed in
the Appendix and discussed in detail in Chapter III.)
-
- Unknown to us, the Italians had already
succeeded in obtaining airfoil force data through the
supercritical range up to about Mach 0.94, and the Germans to
about 0.92. We first learned of this in 1944-before any of the new
Langley schemes materialized-upon the arrival of Antonio Ferri,
formerly of the Italian aeronautical laboratory at Guidonia, and
recently an Italian Partisan in the war. Ferri brought with him
extensive airfoil data from their tests in a semi-open high-speed
tunnel in the early forties. He completed analysis of the data at
Langley [38] and we published the results in a NACA wartime
report (ref.
45). His English at that time was
negligible, and I wrote the final text after much heated consultation
with Ferri and help from Lou Nucci who acted as interpreter.
(Major confusions arose from Ferri's pronounciation of "subsonic"
and "supersonic," both of which sounded to me like "soupsonic.")
The proportions of Ferri's tunnel (1.31 feet across the open top
and bottom and 1.74 feet on the closed sides) corresponded to 43
percent of the perimeter being open. This closely approached the
value of 46 percent suggested by a theoretical analysis of
Wieselberger in Germany (ref. 46) as the correct proportion for zero "blockage"
(zero axial-velocity correction) applicable to a three-dimensional
test model. However, this large degree of "openness" had a serious
drawback in the large pulsations which occurred at high speeds.
None of us was quite sure of the validity of the semi-open tunnel
technique at that time.
-
- Despite the questions of technique,
Ferri's data revealed a most important new finding: the loss in
lift associated with the compressibility burble did not persist
indefinitely. At about Mach 0.9 a marked recovery in lift
occurred, suggesting that the separated ("shock-stalled") flow
tended to disappear as Mach 1 was approached. Later that year
support for this result was indicated in tests of small wings by
means of the "wing-flow" technique (ref. 47). In 1946 we obtained German airfoil data from
their large 2.7-meter closed-throat tunnel (ref. 48) which provided further verification at speeds up
to Mach 0.92. And early in 1947 the first airfoil pressure
distributions ever obtained at Mach 1 were successfully measured
in our rotating-disc annular transonic tunnel (ref. 49). These showed conclusively that at Mach 1 the
shocks had moved to the trailing edge and the flow was supersonic
about the entire section except for a small region at the blunt
leading edge. The German tests had included a systematic study of
the effects of airfoil camber at high speeds which clearly showed
that conventional positive camber was undesirable for Mach numbers
greater than 0.75, and in fact best lift-drag (1/d) ratio was
obtained with negative camber at supercritical speeds, a result
with which Ferri's data agreed.
-
- I had become involved in a study of all
available transonic data in 1947 in connection with writing a
chapter on "Transonic Aerodynamics" for a prospective aeronautical
handbook (ref. 50). At Stack's suggestion
[39] I discussed my
airfoil material at a meeting of the Langley
General Aerodynamics Committee on January 16, 1948. This was
the first time that many of the members had seen the German
results and the general agreement of all of the new data as
regards flow phenomena and trends of airfoil performance at
supercritical speeds approaching Mach 1 (ref. 51). Dick Whitcomb was an interested participant at
this meeting. In commenting on the effects of camber at
supercritical speeds Whitcomb suggested that upper surface
curvature might be the important parameter and that the use of
"proper" curvature might reduce the upper-surface shock strength
and tendency of the flow to separate (ref. 51). Some 16 years later he would resurrect this idea
and apply it successfully in the "supercritical" airfoil (see
pp. 55ff.).
-
- A few months after this meeting I
presented the unclassified parts of my summary material at a NACA
University Conference (ref. 52). Airfoil shock and separation patterns inferred
from the available force and pressure data (refs. 47, 49) throughout the transonic zone were illustrated
(fig. 3). The points brought out in the discussion
included:
-
- The region of shock-stalled flows
"compressibility burble") was limited to speeds approximately
between Mach 0.75 and 0.95.
- By Mach 0.93 in most cases the shock had
moved off the rear of the airfoil, the flow field was entirely
supersonic except for a small region near the leading edge, and no
significant viscous separation effects were present.
- Beyond Mach 1 a smooth transition to
purely supersonic airfoil characteristics could be
expected.
- Camber was undesirable beyond about Mach
0.75. (Schlieren pictures were used in ref. 52 to contrast the shock effects on a thin uncambered
airfoil of low surface curvature and the much larger adverse
effects on a cambered airfoil of large surface curvature.)
-
- Ferri's successful use of the semi-open
tunnel, together with encouraging results of Langley studies of
this configuration by Donaldson and Wright (ref. 53) and Lindsey and Bates (ref. 54) led to our decision in the fall of 1947 to convert
the 4 x 18-inch high-speed tunnel to the semi-open arrangement
with the object of systematic airfoil testing at Mach numbers up
to 0.95, and higher if possible. The first results, obtained in
1948, showed that the tunnel (now 4 x 19 inches in size)
[40]
could be operated with 4-inch chord models at a nominal Mach
number (if 1.0, but it was not immediately certain that the sonic
results were valid. This tunnel was ideally proportioned for
schlieren photography, and from the start impressive photographs
were obtained which provided the first visual proof that our
speculations about the flows at Mach 1 based on force and pressure
data were correct. Figure 4, constructed from photographs taken in
1949, contains typical results from this program. The top row of
photographs, for Mach 1, are of particular interest, showing that
the shocks lie downstream behind the trailing edge. The flow on
the airfoil is virtually separation-free and entirely supersonic
in character except for a small subsonic region near the leading
edge. Extensive systematic pictures of this kind for other
sections were obtained in the 4 x 19-inch tunnel by B. N. Daley
and R. S. Dick and published later (ref. 55). Similar flow pictures were also obtained....
-
-
-
-
- FIGURE 3-Airfoil flow patterns at
transonic speeds discussed at NACA University Conference, 1948.
-
-
[41] FIGURE 4.-Transonic flows and pressure distributions, Mach.
0.79 to 1.00. Angle of attack, 3.2 deg. From the 4 x 19-Inch
Semi-Open High-Speed Tunnel, 1949.
-
-
- [42] ....by the
British some years Langley work and by were used by W. S. Farren
in his Wilbur Wright Memorial Lecture of 1955 (ref. 14).
-
- It is most important to note that our
large burst of understanding about airfoil behavior beyond shock
stall was acquired in the 1945-1947 time period, several years
before any data on this problem were obtained from the research
airplanes. My summary airfoil paper (ref. 52) was prepared in the spring of 1948, before the X-1
pressure data had been
obtained. When I first saw the X-1
pressure data for the 10 percent thick wing about a year later the
fact that it confirmed what was already known from the wind
tunnels was satisfying but not at all surprising. Nevertheless the
wing pressure distributions obtained in flight on the X-1 were of
very great value because they provided the ultimate indisputable
basis for judging the relative merits of the various ground
facilities. The basic airfoil used on the number 2 airplane was
the NACA 65-110, and both the Annular Transonic Tunnel and the 4 x
19-Inch Semi-Open High-Speed Tunnel programs had scheduled this
section for their initial tests in anticipation of critically
important comparisons. (A minor flaw in the plan was discovered
after the tests had been made; in building the airplane the slight
cusp in the basic 65-110 section had been removed for structural
reasons, and this caused a minor change in the flight pressures
just ahead of the trailing edge.) Figure 5 compares the X-1 flight data with the results from
the two transonic facilities at Mach 1. Agreement with the 4 x
19-inch tunnel was considered excellent. The annular transonic
tunnel data, although showing the generally correct shape,
indicate pressures consistently too high. This same type of
discrepancy was noted in subsequent tests of other sections and
was never satisfactorily explained. Of the several transonic
techniques only the 4 x 19-inch semi-open tunnel remained active
in airfoil testing throughout the concluding years of the NACA
program.
-
- An early airing of our new knowledge of
airfoil behavior near Mach 1
was made by Daley and Habel at the
NACA Transonic Airplane Design Conference of September 1949
(ref.
56). During preparations for this
meeting both the 4 x 19-inch tunnel data and the X-1 data were so
new that Daley balked at presenting them without more time for
analysis, but he finally yielded to management pressure. No
conclusions were drawn, however, concerning the relative merits of
the test techniques.
-
-
-
-
- [43] FIGURE 5-The X-1
pressure distributions compared with those of the Annular
Transonic and the 4 x 19-Inch tunnels. Mach 1 .0, NACA 65-110
airfoil, c. = 0.41.
-
-
- The discovery that airfoil flows beyond
about Mach 0.95 did not suffer from significant viscous separation
effects lent new encouragement to the theorists. It had been
previously believed almost universally that sonic flows in real
gases would be characterized by large viscous separation effects,
so that any theoretical treatment, based on the usual ideal gas
assumptions, would have little realism. Thus the main theoretical
concentration up to the mid-forties was on refinement of
subcritical compressible-flow calculations. While this was
consistent with the original belief that practical aircraft would
not be able to operate much above the critical speed, in
retrospect it is apparent that these efforts were not
[44]
very profitable. The simple approximations developed early in this
work were adequate for most engineering purposes, although minor
refinements were laboriously attained later.
-
- By 1947, however, an increasing number of
theorists, encouraged by the new experimental findings, started to
tackle the transonic problem in a variety of new ways. The
development of the transonic similarity laws was a useful aid in
data correlations, although these laws, of course, provided no
solutions for any flow problems. Major progress came when the
special case of the wedge airfoil at zero angle of attack at Mach
1 was solved by Guderley and Yoshihara in 1948 (ref. 57). I was privileged to see this accomplishment
before its publication when Guderley visited Langley to discuss
the work with A. Busemann, who had been assigned to the
Compressibility Research Division after he had been brought to
this country under the auspices of a Navy postwar program. For
some years previous Busemann had consulted with Guderley on this
problem and had contributed suggestions for its solution. The
initial solution was for a cusped wedge shape, but this was
followed shortly by similar results for a symmetrical double wedge
(ref.
58). These results were very
important to us; at long last we had theoretical sonic pressure
distributions against which the experimental data from our new
test techniques could be evaluated. These assessments would
constitute a valuable supplement to the X-1 data as a means of
insuring the validity of the experiments. I enthusiastically
arranged for tests of the wedges in both the Annular Transonic
Tunnel and the 4 x 19-Inch Semi-Open High-Speed Tunnel
(ref.
59). The gratifying results
(fig. 6) were presented for the first time at the September
1949 conference (ref. 56). A photograph of the flow about the wedge at Mach
1 confirmed the absence of any significant viscous separation
effects except for a very small bubble just downstream of the
sharp crest.
-
- An important feature of the Guderley flow
field was a region of smooth shockless deceleration of the local
flow downstream of the crest of the wedge, caused by reflections
of the expansion waves from the curved sonic line extending upward
from the crest. The reflections from this free boundary were
compression waves which decelerated the flow in smooth reversible
fashion. Previously, for conventional airfoil shapes at low
supercritical speeds, no such shockless compressions had been
[45]
identified and it was thought that shocks were the inevitable
device employed by nature to return the flow of stream
velocity.
-
- We now know that a considerable degree of
smooth recompression prior to the terminal shock can occur for a
wide variety of airfoil shapes near Mach 1. Actually this could be
seen in the conventional airfoil pressure distributions obtained
in the late forties at Mach 1 (see fig. 4, top row, for example). A more direct indication of
the effect can be seen in fig. 7 which shows a Mach 1 pressure distribution obtained
in 1948 in the Annular Transonic Tunnel. By comparison with
supersonic expansion theory the measured pressures over the
rearward portion of the airfoil were unaccountably high, and not
understanding the possibility of the recompression effect we
speculated that boundary layer growth might be the cause. The
effect is actually primarily recompression due to reflections from
the sonic fine and secondarily the boundary layer contribution.
Theoretical treatment revealing that all conventional....
-
-
-
-
- FIGURE 6. Guderley theory for Mach 1
compared with Langley transonictunnel data.
-
-
-
- [46] FIGURE
7.-Pressure distribution obtained in the Langley Annular Transonic
Tunnel at Mach 1. NACA 66-006 airfoil at zero angle of attack.
-
- .....sections experience this
effect near Mach 1 came about 10 years later in 1959
(ref.
60). Of great interest here is the
implication that smooth recompression can in some circumstances
also play a major role at speeds well below Mach 1 in the
achievement of improved supercritical airfoils, accomplishing the
benefit suggested by von Karman in 1941 (ref. 61).
-
- One of Busemann's first projects after his
arrival was to summarize the theoretical possibilities for
treating transonic flows, starting at Mach 1 and extending upward
in speed through the detached shock region (ref. 62). Applying these methods, Vincenti and Wagoner
extended the flow field calculations for the wedge to low
supersonic speeds with detached bow waves, showing that the
transition to pure supersonic flow with attached shock was a
stable, orderly process (ref. 63). These results tended to support the conclusion we
had already come to from [47] the experimental
work, that there was little need for systematic experimental
airfoil research in the supersonic part of the transonic region.
We believed that such wing investigations as would routinely be
made in the course of configuration development in the slotted
tunnels and supersonic tunnels would be sufficient, and later
experience proved this assessment correct.
-
- The development of airfoils with improved
supercritical characteristics was a major thrust of the 1945-1955
decade. Nearly everyone working in this field naturally thought of
the possibilities of achieving a "delayed compressibility burble"
Stack in 1938 (ref. 28) and von Karman in 1941 (ref. 61) specifically discussed this possibility. The term
"supercritical" in its broadest sense means any speed beyond the
critical Mach number, but as used by most of us in that period it
meant speeds greater than the force-break speeds and extending
upward into the sonic or low supersonic region. In recent years
Whitcomb has introduced a more restrictive meaning: his
"supercritical" airfoil is designed to delay the drag rise and
thus the term refers to airfoil operation in the speed region
between critical Mach number and drag-rise Mach number.
-
- In a sense, the "dive-recovery" flaps
developed for the P-38 were the first attempt to obtain an airfoil
with improved supercritical performance (ref. 20). Throughout the forties, the tendency of diving
aircraft to lock into a severe nose-heavy condition from which
recovery was often difficult remained the principal problem for
supercritical research. The buffeting which accompanied the lift
loss in shock-stalled flows was a parallel concern. It had become
generally accepted by the mid-forties that high critical Mach
number was no index of good supercritical performance. There is
little correlation between critical Mach number and force-break
Mach number for a wide variety of sections. It was generally
agreed that new criteria would have to be found for the design of
airfoils with good supercritical performance. H. J. Allen came up
with a fresh idea for minimizing the lift loss and moment changes
at shock stall, which was tested with some success (ref. 64). He reasoned that if both upper and lower surface
flows reached local sonic velocity at the same flight speed, a
more equal separation would occur on each surface, leaving the net
lift relatively unchanged. He and D. Graham developed an airfoil
having an "M-shaped" camber line which achieved
[48]
a reasonable approximation of this type of flow. Unfortunately, it
had high subcritical drag and was never used as far as I have been
able to learn. Nevertheless, it was the first attempt to tailor a
specific fixed-geometry airfoil for alleviation of the shock-stall
lift loss; no mention was made of improvement in supercritical
1/d.
-
- In the early forties when the P-38 was in
trouble, I recall a conversation with Allen and Stack in which we
agreed that conventional cambered airfoils showed improved
supercritical lift and moment performance if operated inverted in
the negative-lift attitude (i.e., with negative camber in
the positive lifting sense). Negative camber meant a less curved
upper surface which had reduced separation losses at shock stall.
Allen dismissed this approach as being rather unthinkable and
remarked facetiously that pilots would hardly accept inverted
flight as a technique for pulling out of supercritical dives. None
of us gave much thought to the supercritical lift-drag ratio at
that time; I was certainly unaware that negative camber in
addition to the lift-loss
benefit resulted in better
supercritical 1/d until I noticed that this was so in editing
Ferri's airfoil report in 1945. I looked back at our own data and
some 1945 Ames data (ref. 65) obtained in systematic airfoil tests in their 1 x
3.5-foot tunnel at speeds up to about Mach 0.85, and noted with
interest that the supercritical 1/d was significantly better for
negative camber. Figure 8 taken from the Ames data shows this
result.
-
- My past upbringing to the effect that
positive camber was inherently
beneficial and essential to
conventional lifting airfoils at normal speeds was so deeply
ingrained that I dismissed these results as an impractical
aerodynamic curiosity. Three years later, in 1948, I included in
my summary NACA Conference paper on high-speed airfoils
(ref.
52) a plot based on the German
airfoil data (ref. 66) which showed in detail how the camber for best l/d
quickly diminished to negative values as the Mach number advanced
beyond about 0.75 (fig.
9). Actually the data clearly
showed that negative
camber (dashed lines on
fig. 9) gave best l/d at the higher speeds. But still
believing that negative camber was unthinkable for practical
applications, I terminated the plots at zero camber and suggested
as a major conclusion that zero camber (symmetrical) airfoils were
the best compromise for transonic applications (ref. 52). This interpretation was shared by the other
airfoil.....
-
-
-
-
- [49] FIGURE
8.-Improvement in supercritical lift-drag ratio (1/d) with
negative camber. Ames data,1945.
-
-
- ....specialists at Langley at that time
and also by Allen and Graham (ref. 64).
-
- The physical mechanism by which the
improvements due to negative
camber came about was thought at
that time to be related to the location of the peak suction
pressure and shock near the leading edge for the "peaky"
distribution of pressure that occurred over the relatively flat
upper surface with negative camber. For this forward shock
position, the boundary layer was thin and not as prone to
separation as it was for the positively-cambered case where the
shock occurred far aft on the curved afterbody where it triggered
separation. We did not realize then that an additional mechanism
was at work for the "peaky" case, namely some degree of shockless
recompression due to reflections from [50] the sonic line.
("Peaky" is a term coined a decade later by Pearcey, who called
attention to the recompression effect (ref. 67).)
-
- The first attempt to derive
"supercritical" airfoils in the restricted Whitcomb sense was made
in 1951 by Woersching (ref. 68). He had studied all of the available negative
camber data including the negative lift operation of the
Allen/Graham "M" cambered NACA airfoil 847B-110 (ref. 64). He noticed that this airfoil in the inverted or
negative lift attitude had a drag-rise Mach number of 0.81 at
c1
= 42, while in the normal attitude the drag rise occurred at M=
0.73. He also examined the inverted airfoil in the region of shock
stall and beyond and found it to be generally as good as or
superior to the normal attitude. After further study, Woersching
concluded, "Maximum drag rise Mach number is obtained with
negative camber over the forward chordwise portion of the airfoil,
and positive camber aft to the trailing edge-but at the expense of
large negative moment coefficients." This, of course, is a
qualitative description of the features of the Whitcomb
"supercritical airfoil"-together with one of its special problems.
Woersching goes on to advocate inclusion of the last arm of the
"W" camber in order to relieve the pitching moment problem at some
loss in drag-rise Mach number. He also visualized aircraft
incorporating both sweep and the proposed sections, designed "for
cruise near Mach 1." This work was
undoubtedly the first serious
attempt at delaying the drag rise-with a profile that would
qualify as a "supercritical airfoil" in the present-day sense. By
way of explanation of the action of negative camber Woersching
pointed out that it results in a degree of flatness of the suction
surface comparable to that of a much thinner symmetrical section.
Unfortunately, he did not have the resources to continue
development.
-
- Probably inspired by the Woersching paper,
Britisher W. F. Hilton published in 1953 a report (ref. 69) which he had written in 1947 on the advantages of
negative camber. The original report had apparently been given
only restricted circulation in Great Britain, perhaps for reasons
of security. It is interesting to note that Hilton had been
employed in the United States for several years following the war
and had access to and personal interest in the available American,
Italian, and German airfoil data. Hilton did not recommend any
particular.....
-
-
-
-
- [51] FIGURE 9.-Effect
of Mach number on camber for best lift-drag ratio.
12-percent-thick airfoils.
-
-
- ....distribution of negative camber. His
aim was primarily to reduce the adverse lift and moment changes
due to shock stall and secondarily to improve I/d beyond shock
stall.
-
- Without doubt the period from 1945 to 1951
was one of the most productive eras in the history of high-speed
airfoil research. Several new transonic ground facilities and
flight techniques were developed and applied successfully;
reliable wind tunnel data at Mach numbers up to 1.0 were obtained,
including airfoil flow photographs; new theoretical treatments of
the flow were accomplished for wedge airfoils at Mach 1 and
throughout the detached shock range; criteria were established for
airfoils having delayed drag rise and an inverted NACA airfoil
meeting these criteria was specified (the first "supercritical"
airfoil in today's [52] parlance). NACA
program activities were at the core of this progress, although
there were also important outside contributions, especially on the
theoretical side. On June 2, 1950, 1 reviewed this satisfying
progress in considerable detail for the NACA Executive Committee,
concluding, "The principal details of two-dimensional transonic
flow are now known as a consequence of recent progress both
experimental and theoretical.... Many problems remain for the
three-dimensional case of complete wings. . . . Our 8-foot
high-speed tunnel with its new slotted throat provides a transonic
facility of adequate size for the needed work on complete
wings."
-
- John Stack's role in the high-speed
airfoil developments of this period was quite different than his
intimate personal participation during the first dozen years of
his career. From about the time of his Wright Brothers lecture it
had seemed likely that he would be moved into a management
position in the Langley "front office." His special talents as a
tough, persuasive technical salesman were badly needed and,
furthermore, it was obvious that life would be much more pleasant
for Langley management with Stack as a member of their office
rather than as a combative division chief who increasingly was
cast in an adversary relationship to higher management in regard
to approvals and funding allocations for our projects. Thus in
mid-1947 Stack became an assistant to Chief of Research F. L.
Thompson, and I succeeded him as Chief of the Compressibility
Division.
-
- Although he remained invariably supportive
of our projects, my relationship with Stack was inevitably
changed. He was now one of "them" rather than a close colleague in
research. His principal preoccupation became the promotion and
development of major new transonic and supersonic tunnels, and he
also became involved with other problems beyond our field of
interest. He observed the airfoil developments with interest as
they unfolded but had no direct part in them except through
related facility developments-such as the Annular Transonic
Tunnel, which might never have been successfully promoted without
Stack's support. His early experiences with the open-throat
tunnels made him rather suspicious of semi-open tunnels and this
was reflected in his encouragement of studies of their transient
disturbances by Lindsey and Bates (ref. 54). In contrast to Stack's many publications in his
earlier [53] airfoil research period, the paper covering his
review of facility developments (ref. 54) was his principal publication in the 1945-1951
period.
-
- In the final period of the NACA program
from 1951 through 1956 a rapid dwindling of the effort took place.
This was due partly to the large shift in research emphasis to
swept and low-aspect-ratio wings for supersonic aircraft, and
partly to the fact that a substantial high-speed airfoil
technology base had been established. The demand for
two-dimensional airfoil research diminished to low levels except
for the special area of helicopter blading. Ames experimental work
in the field V4, had been terminated in 1951 when their 1 x
3.5-foot tunnel was phased out as a closed-throat facility.
Finally, the abolishment of W. F. Lindsey's section at Langley in
1956 brought to a close the NACA high-speed airfoil program which
had started 29 years before. Although several worthwhile projects
were left unfinished, they could not compete in priority with the
demands of supersonic aircraft and the burgeoning space
program.
-
- COMMENTARY
-
- Curiously, the impressive progress
in-high-speed airfoil technology in the last decade of the program
is often overlooked. At a recent NASA airfoil conference
(ref.
70) several practitioners in the
current program seemed to believe that the NACA program had
terminated with the 16-series sections and Stack's Wright Brothers
Lecture of 1944 which, so to speak, left high-speed airfoils in
the depths of the shock staff. The most likely explanation is that
the researchers of 1945-1956 did not produce any specific new
airfoil families. They produced important new understanding of
transonic flows and they extended the accurate data for existing
airfoil families to Mach 1, but unfortunately, perhaps, there were
no associated clever baptisms or new acronyms to help publicize
the progress that was made. This solid but unspectacular airfoil
progress was overshadowed by the more dramatic events of that
period-the first supersonic flight, the slotted transonic tunnel,
and the area rule.
-
- The high-speed airfoil program provides an
excellent example of NACA accomplishing its mission in an
important problem area of aeronautics. For the first 20 years,
from the early twenties to the early [54] forties when the
propeller was the primary application, the program provided both
fundamental understanding of the flow phenomena and new airfoils
to improve propeller performance at high speeds. These solutions
were in hand well before they were needed by industry. Only for
the brief period from about 1944 to 1947 was the program deficient
in meeting the new needs for transonic data beyond Mach 0.85.
During this interim, foreign data plus information from the
"wing-flow" and "body-drop" techniques were used effectively.
Early in 1947 the first airfoil/pressure data at Mach 1 were
obtained (ref. 49) and by 1949 an effective semi-open wind tunnel was
being used routinely for airfoil testing to Mach 1, the technique
verified by X-1 flight data.
-
- The program wound down rapidly in the
mid-fifties, partly because there was no obvious need then to
expand the technology beyond its already substantial proportions,
but mainly because several of its talented researchers had been
lured into more urgent and fascinating supersonic and
space-related projects. Almost a decade would pass before the
renaissance described in the next section would take place, based
on the recurrence of an old need, but carried forward with fresh
inspiration by a wholly new research team.
-
- The report editing procedure mentioned in
this chapter and elsewhere deserves comment. The primary technical
editing was accomplished by an inter-divisional committee of the
author's peers. This was followed by editing for grammar,
availability of references, etc., by a female "English critic" in
the editorial office. The generally superior reliability, clarity,
and freedom from "governmentese" of the NACA reports produced by
this system have been widely acclaimed. Unfortunately, however,
most of them are rather dull from a literary point of view.
The report-writing manual
used to indoctrinate young NACA
engineers emphasized accuracy, clarity, and adherence to the
standard format, rather than any matters of style or technique to
make the report
interesting. Language which added
humor or sparkle was frowned on and almost always deleted.
Imaginative speculation was forbidden unless specifically
identified as such. All of this was perhaps appropriate for simple
reports intended to present reliable data in a readily usable
form. But by the time NACA writers had progressed to more
sophisticated subjects such as advanced concepts or
state-of-the-art papers for a [55] national
audience, most of us were so crippled by habitual adherence to the
system that these writings also tended to be stereotyped and less
interesting than they might have been.
-
-
-