section in the closed return circuit, and a low-speed (135 knots)
.', section in the closed return circuit, and a low-speed (135 knots) 130x200 ft. test section in an open return circuit. The third configuration is an open return wind tunnel with a test section size of 75x150 ft. and a speed of 300 knots. The open-return design is of interest because it minimizes struc~ural cost, which is the major cost factor in facilities of this size. All of these facilities have power requirements of about 425,000 horsepower.
ECONOMIC CONSIDERATIONS The overall judgement that must be made in considering the proposed facility is whether the research and development value of the work performed in this facility will justify its cost. Since this facility should have an operational life of at least 25 years, it is obviously impossible to make a detailed and specific cost-benefit analysis.
However, it is possible to arrive at a reasonable perspective on the value of the proposed facility by considering three factors: <a) the alternative of flight tests; (b) the estimated cost of future aircraft programs and the possible contributions of the full-scale wind tunnel in reducing these costs; and (c) the experience of ·theexisting full- scale wind tunnels. .
Flight Tests as an Alternative to Full-Scale Wind Tunnel Tests " Since many of the requirements for the full-scale wind tunnel presented in previous sections of this report dictate tests of full- scale hardware, the merits of full-scale wind tunnel tests vs. flight I w., tests were examined. There are several reasons why full-scale wind tunnel testing may be necessary or advantageous. For example, rotor dynamic stability or engine stall problems involve safety-of-flight, and the loss of an aircraft due to these causes generally results in the grounding of the flight test aircraft. Under these circumstances further flight tests are not possible. Another example where flight testing may be quite hazardous is during flight operation close to the ground, where the high flow distortions described prev~ously in connection with engine stall may be encountered. An engine stall or loss of control during such. conditions might easily lead to loss of the aircraft, since very little altitude is available to recover from a ; .....
severe upset.
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Aside from the risks involved in flight tests, the quality, quantity, and type of test data that can be obtained in flight is limited by lack of control over the test environment, cost, and the constraints due to the fact that controlled flight must be maintained. For example, tests to identify the specific contributions of various aircraft components by progressive tests with various components removed are not possible in flight tests, but are a valuable standard technique in the wind tunnel.
Again, the ability to vary one significant parameter while holding all other parameters constant is a valuable wind tunnel test technique which cannot be done in flight. In Ref. 11 several comparisons are made between flight tests and tests in ground-based facilities. Among the conclusions are: . (a) The number of measurement channels available is generally less in flight testing.
(b) The accuracy of flight test data is usually about 5% compared to accuracies of 1-2% in ground-based facilities.
(c) The cost of flight tests are an order of magnitude more expensive than those for tests in ground-based facilities.
Costs of Typical Aircraft Programs The prime'justification for the proposed facility is that the cost of the facility is reasonable in terms of the cost of the aircraft programs which it would support, and in terms of the savings that can be realized for these programs over the costs that would be incurred in the absence of the facility. This assessment is difficult to make in specific cases, since the operational life of the proposed facility would be at least 25 years (i.e. from about 1977 to 2002). However, some perspective on this can be realized by considering the costs of typical aircraft programs such as are described in Ref. 12.
Figure 12 compares the fUnding schedule for a typical aircraft pro- gram (from Ref. 12) with the cost of the proposed facility. Figure 12 shows that the first major input from tests in the facility occurs prior to flight test when both the level of funding and the rate of increase of funding are low. These tests typically would involve full-scale demonstrator hardware of the high-risk items (e.g. rotors, fans, engines) installed in inexpensive "boiler plate" airframe mockups. Data obtained at this early stage of the program has extremely high leverage on program costs, since the funds committed are still low. The second period of major influence of the facility shown on Fig. 12 corresponds to the early flight test stage. ·If unanticipated problems are encountered in flight, '0 the aircraft can be returned to the full-scale wind tunnel for rapid and safe exploration of these problems. The alternative to this is continued flight tests with the accompanying risks to the aircraft and pilot. By the time flight tests can resolve major discrepancies the cost of the program exceeds that of the facility. In addition, by that time commit- ments have been made such that the cost incurred by changing the design may be as much as the funds expended. Thus, the cost of the proposed facility should be viewed as an insurance premium which reduces the risk of potential losses in advanced technology programs to acceptable levels.
It is quite likely that, without the assurance provided by early tests of critical components in the proposed facility, many advanced aircraft programs (e.g., high-speed rotorcraft and V/STOL aircraft) will not be initiated due to excessive financial risk.
Experience with Existing Full-Scale Wind Tunnels NASA has operated full-scale wind tunnels since the late 1920's, and there is considerable historical evidence of the value of this type of facility. The first NASA full-scale wind tunnel (the 20-ft. I ~ •• Propeller Research Wind Tunnel) showed the absolute necessity of using cowls for radial engines and variable pitch propellers for high performance aircraft. The incorporation of these features provided the first high performance transports which launched the era of practical . \ commercial air transportation. Similarly, the application of this research to fighter aircraft culminated in the high-performance radial engine fighters of World War II fame.
The prime justification for construction of both the 30 by 60 and the 40- by 80-foot wind tunnels was for the performance of drag reduction studies. Drag reduction work in the 30- by 60-foot wind tunnel just prior to, and during the early part of World War II, added an average of 30 MPH to our World War II fighters. The sources of drag reduction were each small in themselves, but the total effect was large. They included such items as unnecessary air leakage, unfaired scoops, improper cooling air discharge, necessity of using cowl flaps to control cooling airflow, etc. This drag reduction and cooling work had to be done with the full- scale hardware since it involved details of the interface between construction practices and details with aerodynamics.
While the prime justification for both the 30- by 60 and the 40- by 80-foot wind tunnels was for drag reduction studies of military air- craft, the major contributions of these facilities have been in technological areas which were not anticipated at the time the facilities were planned. The contributions of these facilities to the research and development of V/STOL aircraft is an example of this. The'40- by SO-foot ~... .
, wind tunnel has more than paid for itself by preventing catastrophic failures of experimental V/STOL aircraft in flight. These aircraft encountered failures during tests in the 40- by aO-foot wind tunnel which could have been catastrophic in flight. All of these failures involved the complicated interface between aerodynamics, dynamics, and structures. Therefore, tests of the full-scale hardware were the only ,. " way that these problems could have been discovered. For example, the XV-I compound helicopter encountered a rotor speed instability during the wind tunnel tests which required changes to the rotor control system. Tests of the XV-3 in the 40- by aO-foot wind tunnel were requested after a catastrophic rotor-pylon whirl instability was encountered in flight, resulting in loss of the aircraft and serious injury to the pilot. After two tests in the 40- by aO-foot wind tunnel, separated by a one-year analysis effort, this stability problem was alleviated so that a highly successful flight research program could be completed. As a result, the tilt rotor aircraft is considered today to be one of the more promising high performance rotary wing aircraft concepts. The first wind tunnel test of the XH-51 rigid-rotor helicopter ended in a catastrophic break-up of the rotor due to a bonding failure, The rotor blade was redesigned, a successful wind tunnel test was completed, and the XH-5I went on to a highly successful flight research program which culminated in a rotorcraft speed record. During wind tunnel tests of the XV5A lift-fan airplane, structural failure of the fan inlet guide vanes was encountered. If these had failed in flight and entered the fan rotor the aircraft would have been lost. Also, excessive deflection of the fan exit louver control mechanism was encountered during these wind tunnel tests. This would have severely limited the fan-supported flight envelope of the XV-5A. Both of these problems were remedied following the wind tunnel tests, and the XV-5A airplane has completed a series of successful flight research programs.
The lift-fan propulsion system is currently considered to be one of the most promising concepts for a high performance V/STOL airplane.
The total cost of the aircraft programs -which have been saved by tests of the full-scale aircraft in the 40- by SO-foot wind tunnel has ,~ more than offset the total cost of construction and operation of this facility for 25 years. To these savings could be added the savings due to the cancellation of flight test programs of aircraft which had been shown by full-scale wind tunnel tests to have fundamental deficien- cies. A partial list of such aircraft is the Kaman K-16 tilt wing, the Avrocar, and the Vanguard low-disk loading fan-in-wing airplane.
Actually, the most important contributions of the full-scale wind tunnels have been in research areas where it is nearly impossible to put a firm dollar magnitude on the value of the contribution. The contributions of the full-scale wind tunnels to the development of the externally-blown flap and the augmentor wing turbofan STOL aircraft are ; ~ . , almost solely responsible for the fact that these concepts are today :~ I considered to be the most promising types for application to large commercial and mi~itary STOL transport aircraft. The use of the fu1l- scale wind tunnels was also instrumental in establishing the feasibility of conventional landings of lifting body spacecraft. This included , .
tests of all of the full-scale flight vehicles in the 40- by 80-foot wind tunnel plus studies of free-flight models in the 30- by 60-foot wind tunnel. These studies added immeasurably to the confidence level required", before flight tests could be initiated with these radically new and ...
different aircraft. This is perhaps the best example of an application of the full-scale wind tunnels to fill a need which could not have been anticipated when the facilities were justified.
Since the cost of the proposed full-scale wind tunnel has already been compared with the typical cost of current aircraft programs •. it is.
of interest to compare the cost of the existing full-scale wind tunnels with representative aircraft current at the time these facilities were justified. This comparison is presented in the following table.
TABLE I Cost of Full-Scale Wind Tunnels Relative to Aircraft
Year Cost of Alc Cost of F.S.W.T. 'W.T. Cost
~/C Cost $25,000 $1,000,000 (30x60) 40 1940 130,000 7,000,000 (40x80) 53 1975 7,000,000 200,000,000 28 This comparison shows that the cost of the proposed full-scale ( wind tunnel relative to the cost of current aircraft is of the same order as the comparable relative costs for the 30- by 60- and 40- by 80- foot wind tunnels. In addition, the need and the potential for both military and civil air transpor~ation is much more apparent now than it was at the time the existing full-scale wind tunnels were justified.
In retrospect our predecessors showed a high degree of foresight and courage in building the 30- by 60-foot wind tunnel and 40- by 80-foot wind tunnel during the great depression and at the outset of World War II, respectively. It is now our turn to make a similar investment in facilities to ensure our future in the aeronautical world.
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, • REFERENCES 1. Anon.: Report of the Joint DOT-NASA Civil Aviation Research and Development Policy Study, Vol. II, NASA SP-266, 1971.
2. Wimpress, J. K.: Aerodynamic Technology Applied to Takeoff and Landing. Annals of the New York Academy of Sciences, Vol. 154,' Art. 2, Pages 962 to 981, Nov. 1968.
3. Deckert, W. H.; and Hickey, D. H.: Sununary and Analysis of Feasibility-Study Designs of V/STOL Transport Aircraft, J. Aircraft, Vol. 7, No.1, Jan.-Feb., 1970.
4. The Boeing Company: Study of Aircraft in Short-Haul Transportation Systems, NASA CR-986, 1967.
5. Fry, B. L.; and Zabinsky, J. M.: Reasibility of V/STOL Concepts for Short-Haul Transport Aircraft, NASA CR-743, 1967.
6. Marsh, K. R.: Study of the Feasibility of V/STOL Concepts for Short-Haul Transport Aircraft, NASA CR-670, 1967.
7. The Lockheed-California Company: Study on the Feasibility of V/STOL Concepts for Short-Haul Transport Aircraft, NASA CR-902, 1967.
8. Anon.: Technical and Economic Evaluation of Aircraft for Intercity Short-Haul Transportation, Federal Aviation Agency-Air Defense System-74, Vols. I, II and III, April 1966.
9.. Kuhn, R. E.; Kelly, M. W.; and Holzhauser, C. A.: Bringing V/STOL's Downtown, Astro. Aero., Sept. 1965.
10. Galbreath, A.; and Warfield, R. M.: Terminal Area Airline Delay Data, 1964-1969, Federal Aviation Administration, Washington, D. C., Sept. 1970.
11. Kamchi, J. S.: National Aerospace Test Facilities, NATO Defense Research Group Seminar, Saint-Louis, France, May 4-7, 1971.
D.: Forecasting Military Aircraft, Space/Aeronautics, 12. Harris, N.
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1968. . I I ..
MAJOR PROBLEM AREAS AND ASSOCIATED TECHNOLOGY
FOR CIVIL AIR TRANSPORTATION
PROBLEM AREAS
• NOISE
o CONGESTION
o SHORT HAUL SYSTEMS
ASSOCIATED TECHNOLOGY
o HIGH-LIFT AND V/STOL AERODYNAMICS
o PROPULSION
o ACOUSTICS
o AVIONICS
Figure 1
EFFECT OF MAXIMUM LIFT ON COMMERCIAL TRANSPORT
ECONOMICS
PROFIT FROM INCREASED LANDING LIFT INCREASED LANDING LIFT INCREASED 133% TOTAL PAYLOAD $/TRIP I,OOolb
$1,000
2.4
50.-
2.0
40~
1.6
BASIC 30~ ~
- CL 1.2
20~ LANDING FLAPS
.8
10 10
o 8 12 16 20 24 °BASIC +5% o BASIC +50/0
CL
CLmax
max
aWING deg
(S3 PROFIT Figure 2
o DOC+IOC
u. S. AIRLINE COST FOR TERMINAL AREA DELAYS
/
200-
/
/
/
160 -
120-
DOLLARS,
MILLIONS
80-
o I I I I I I I I
1964 1965 1966 1967 1968 1969 1970 1971
YEAR
Figure 3
ROTOR VELOCITY DIAGRAM
V
max
40x80 ft
WIND TUNNEL
; nR/-IB,
ft/sec
REVERSE
FLOW REGION
-=:...---L----L...-_---L_----I.~ __ ._...J.I __ .;;.;;L;.... _ ____JI.__ _ __II
o
300 200
V/ -18, knots
Figure 4
EFFECT OF FLOW DISTORTION ON ENGINE PERFORMANCE
J'
F'ERFORMANCE LOSS
IOOr-----
• LARGE THRUST LOSS
PERCENT
• UNSTEADY FLOW
OF UNIFORM
STALL
FLOW • BLADE VIBRATION
PERFORMANCE
• TURBINE OVER TEMPERATURE
o
INCREASING INLET FLOW DISTORTION
Figure 5
SOURCES OF FAN DISTORTION
DURING TRANSITION TO FORWARD FLIGHT
Figure 6
ROTORCRAFT SIZE TRENDS
+- 100
DISK LOADING, psf
~ ..
Z <l
3; 80
SINGLE ROTOR, HISTORICAL
C)
TREND (AHS PAPER 552, 1 )
z
-
~ 60
a:
o
..
a:
~ 40
w
~ <l
-
o 20
a:
o
....
o
I I a:
o 10 20 30 40 50 60 70 80
GROSS WEIGHT, Ibs x 10-
I I I
o 20 40 60 80 PASSENGERS
o 4POO 8,000 12,000 16,000 PAYLOAD
WEIGHT,lbS x 10-
Figure 8
HIGH DISK LOADING V/STOL AIRCRAFT SIZE TRENDS
WING-INDUCED
LIFT
~~~~~- VECTORED THRUST
SPAN, ft
~ 40
\A\
100 120 40 60 80 20 o
GROSS wt,lbs x 10-
I , I I
100 120 PASSENGERS 40 60 80 o
20 24 PAYLOAD,lbsxtQ-3 8 12 16 o
Figure 9
:: 250
WIND TUNNEL WIDTH FOR
..
100 ft SPAN AIRCRAFT
::r: I- ~200 -.J
w
Z
z 150
~ I- ".:.:':' : ..
0::
z 100
<t
cp
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en
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., :£l
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0:: 0 0::
-
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100 400
0 200 300
FLIGHT SPEED, knots
Figure 10
SECTION SPEED HORSEPOWER
TEST SECTION SPEED HORSEPOWER 75xl50ft 300 knots 410,000
•
60xl20ft 350 knots 130x200ft 150 knots 440,000 - - - - --{/..,~;A,4-- - __ 15xl50ft 300 knots 410 000 Figure 11 \ I TYPICAL AIRCRAFT FUNDING SCHEDULE AND DECISION POINTS SHOWING TEST LEVERAGE (CUMULATIVE) PRODUCTION ...
DECISION ~ 2500
~
en I o o(/) I ~-- ~ PERIOD OF MAJOR FLIGHT TEST
~ a 2000
<1:= : INFLUENCE ON DESIGN 0::0 ~-c r O,+- t---~I ~--i PERIOD OF MAJOR WIND TUNNEL TEST 0:: 0 1500 I INFLUENCE ON DESIGN a...
(/) we I >.0 I I ~= 1000 I ...JE I ::> :E I ::> u REPRESENTATIVE COST OF I PROPOSED FACILITY
o
3 6 9 12 15 18 YEARS Figure 12