Experimental Facility and Measurement Technique

The tests were performed in the Princeton University 3 x 4 ft smoke tunnel. A sketch of the tunnel is shown in Fig. 2.1 and also on the back cover. It consists of an inlet and stilling chamber 9 ft high by 12 ft wide containing screens and flow straighteners. The flow straighteners are 3 in square and 12 in long. This section is followed by a 9:1 contraction leading to the test section which is 3 ft high by 4 ft wide. Downstream of the test section the flow is turned 90° and exits through a 50 HP fan. The tunnel speed in the test section is variable from 5 to 70 ft/s.

2.1 Flow Quality

Constant temperature hot-wire anemometry5 (using Dantec model 55M01) was used to determine turbulence levels in the freestream. At all conditions the wire was operated at an overheat of 0.8. The frequency response was optimized using the standard test in which a square wave in voltage is injected at the Wheatstone bridge to simulate an impulse in velocity. The —3 dB point of the response curve was 33 kHz for chord Reynolds numbers of 100k, 200k, and 300k; and 25 kHz for a Reynolds number of 60k. As will be shown shortly, these frequencies are well above the energy-containing frequencies of the turbulence.

A common problem when measuring turbulence levels in low-speed facilities is determining the lowest frequency of interest. Usually, the anemometer signal is high-pass filtered. This procedure reduces the apparent RMS turbulence level by removing low-frequency fluctuations which may be important to boundary layer transition. In this work, however, no high-pass filter was used. Instead, the DC component (the mean) of the anemometer signal was subtracted (“bucked off”) using an operational amplifier of an analog computer. The remaining signal was then amplified to fill the ±10 volt range of the 14-bit analog-to-digital converter and sampled at frequencies from 10 Hz to 10 kHz. By sampling over a range of frequencies, high resolution of the spectra was obtained. In each case the low- pass frequency of the filter was set to somewhat less than the Nyquist sampling frequency to eliminate aliasing errors. A sampling frequency of 100 Hz resolved the high-frequency end of the spectrum and extended down to sufficiently low frequencies.

All spectra presented here were found using a sampling frequency of 100 Hz with the hot wire located 3 in below the center of the tunnel. This location was representative of the turbulence characteristics throughout the central region of the test section. For each run, 9216 points were taken and then broken into ensembles of 1024 to calculate spectra. Spectra at several different Reynolds numbers are shown in Figs. 2.2 (а-d). Power spectral density multiplied by fre­quency is plotted against the logarithm of frequency. In this way, the area under the curve is directly proportional to («rme)2/^oo—the square of the turbulence intensity. As can be clearly seen, the majority of the energy is found at frequen­cies below 1 Hz. If the signal were high-passed at 1 Hz, this contribution to the turbulence would be lost. Perhaps fluctuations at frequencies this low have quasi-steady effects; in any event it is currently unclear what cut-off frequency should be used so both numbers are presented. The unfiltered turbulence levels at various Reynolds numbers are given below. Note that these levels correspond to a very low cut-off frequency of 0.01 Hz due to the sampling interval. Turbu­lence levels are also indicated below for the case of a cut-off frequency of 1 Hz.

RMS Turbulence Intensities


> 0.01 Hz

> 1 Hz













Mean-pressure surveys to determine the uniformity of the freestream were taken in the test section throughout a plane perpendicular to the flow. Less than 4% variation was found in the static pressure and there was no measurable total pressure variation. These surveys indicate a 2% variation in the velocity which was deemed sufficiently uniform.


The primary goal of this work was to design a new group of high-performance airfoils for radio controlled model sailplanes. As can be imagined, this involved numerous preliminary steps from preparing and instrumenting the tunnel to arranging for the models to be built, establishing a design procedure and, of course, solving all the myriad problems that occur in a multi-year project of this size.

In order to establish baseline data, a number of existing airfoil designs were tested first. These were selected primarily by the modeling community and are representative of what is presently used. They ranged from very simple, flat-bottom types, as well as some of the older NACA sections and their close derivatives, to very modern FAI-contest airfoils. Aside from providing the base­line, testing these airfoils allowed us to compare our data with other facilities where the same sections had also been tested.

The flow behavior over an airfoil at high Reynolds numbers—greater, say, than 1-3 million—is well known. The boundary layer is laminar from the lead­ing edge to a point typically near mid-chord where it makes a transition to turbulent flow. This transition, as well as the flow behind it, is generally well behaved. Unlike full-size airplanes, model sailplanes typically operate at chord Reynolds numbers between 50,000 and 500,000, often called the low Reynolds number regime. At these low Reynolds numbers the flow is fundamentally dif­ferent and more complicated than at high Reynolds numbers. The transition process is neither abrupt nor does it usually take place while the boundary layer is attached to the airfoil. Instead the laminar boundary layer separates, that is, it physically detaches from the airfoil surface. The flow then becomes unstable while separated, and makes the transition to turbulent flow in “mid-air.” Only then does the flow reattach to the airfoil. And sometimes, if the laminar sepa­ration point is sufficiently far aft or if the Reynolds number is very low, the flow entirely fails to return to the airfoil surface. In either case large energy losses are associated with this process. This laminar separation, transition to turbulence, and turbulent reattachment enclose a region of recirculating flow aptly called the “laminar separation bubble.” It is this extended transition process that is the principal reason for the degradation in performance at low Reynolds numbers. Efforts towards drag reduction, therefore, largely concentrate on reducing the size and extent of the bubble.

In an effort to understand the flow phenomena at low Reynolds numbers in general (which includes RC sailplanes), there have been numerous theoretical and experimental investigations which have resulted in three major conferences in the past four years1’2,3 and a special AGARD publication4. Because of the difficulty in mathematically modeling the bubble, computational efforts have not been entirely reliable in predicting this complex flow. Nonetheless, considerable progress has been made. In most studies and experiments, the primary emphasis has been on understanding the fundamental mechanisms that drive the bubble. Yet despite the high level of interest in this area, few systematic attempts have been made to apply the growing body of knowledge to the problem of airfoil design. This book discusses a major experimental program that was carried out to do just that, and to do it specifically for RC sailplane airfoils.

All the models were tested in the low-speed, low-turbulence, З x 4 ft smoke tunnel at Princeton University. The following is a list of the 54 different airfoils; several airfoils were duplicated in order to examine the effects of model variabil­ity, and these are indicated by an asterisk (*). The DF – and SD-airfoils are the new designs resulting from this work.






















NACA 0009




NACA 2.5411




NACA 64A010




NACA 6409















Flat Plate






In order to ensure the enthusiasm of the modeling community which built all of the test sections, we tested any airfoil that a builder wanted to supply. As a result of this policy, three things happened. First, a great variety of airfoils was tested spanning virtually the entire range of usefulness to RC sailplanes. Second, some airfoils previously unknown to us offered insights into the airfoil design process. And third, we were able to design new airfoils, have models of them built, and test them in the tunnel, thereby “closing” the design loop. As far as we know, this last step has not been done before on this scale for model aircraft applications.

This book has two major parts: (1) the documentation of the facility and the quality of the data; (2) the results of the tests on over 60 wind tunnel models. The first part is covered in Chapter 2 and extensively documents the experimental methods we used in this work. This part is intended primarily for those active in low Reynolds number airfoil research and may be skipped without loss of continuity by those more interested in the data. To help the non-specialist, the terminology we use is explained in Chapter 3. Chapter 4 briefly covers some of the ideas behind the new airfoil designs. Hopefully, a more comprehensive report will come at a later time. The second major part begins with Chapter 5, which discusses the airfoil polars at length. Airfoil polars and lift plots, tabulated data, coordinates, and other supporting data are given in Chapters 7 through 13. And finally, for those who may wish to contact the authors or obtain additional copies of this book, the addresses are listed in Chapter 14.

4 Airfoils at Low Speeds

The history of this experimental program on low-speed airfoils is extensive. In August 1986, work toward testing model sailplane airfoils in a wind tunnel at Princeton University began on an ambitious scale. The initial plan was to test 30 airfoils: 15 existing airfoils and 15 new airfoils to be designed concurrently with the tests. As news of the project caught the attention of radio control (RC) model soaring enthusiasts, the project grew far beyond the original goals and expectations, thanks to their generosity. When the experimental apparatus was finally dismantled in January 1989, almost two and a half years later, over 60 models were tested and over 130 airfoil polaxs were generated. It is our hope that the results of this work will be valuable to modelers and researchers for many years to come.

A word is in order to explain the role each of us played in this effort. The initial impetus for the project, its organization and day-to-day management, as well as the wind tunnel testing and data reduction were done by Selig and Donovan. They also designed all of the new airfoils except for the DF-series by Fraser.

The custom measurement apparatus was built jointly by Selig and Donovan at Princeton University and by Fraser at Fraser-Volpe Corporation. The digitizing of the models was done at Fraser-Volpe Corporation by Fraser, who also wrote the computer programs for reducing this part of the data. All three of us shared in the writing and editing of this book.

We would also like to mention that everything from the data collection to the writing of this book was done by computer. There is not a single number anywhere in any part of the data that was generated, computed, reduced, copied, averaged, printed, graphed, or manipulated by hand. Aside from the speed and convenience of this approach, the principal advantage is the complete elimination of several types of errors that may otherwise occur.

All of the airfoil polar data is available on 3^ or 5-^ inch IBM compatible diskette from Fraser.

Sincere thanks go to Prof. Smits of the Princeton University Gas Dynamics Laboratory for his enduring support while this extracurricular project began to grow and consume seemingly endless hours of time away from the first two au­thors’ regular thesis research. The gracious support of Prof. Lam and Prof. Cur­tiss of Princeton University, and the helpful discussions of Prof. Maughmer of The Pennsylvania State University are appreciated. Thanks also go to Lou Piz- zarello who provided us with an air conditioner and new air intake filters for the tunnel.

We are indebted to Ray Olsen for his many contributions at times when we needed them most.

For monetary contributions which made possible the purchase of important el­ements essential to this work we thank Rolf Girsberger, H. A. Stokely, Jerry Jack­son, Armin Saxer, Charles Griswald, C. Haverlan, H. J. Rogers, Preben Norholm, Brian Smith, Thomas Yamokoski, and Trey Wood.

The expertise of many skilled model builders made the lengthy set-up stage all worthwhile. In this respect we very deeply appreciate the work of Bob Champine, Ron Wagner, Stan Watson, Mark Allen, Michael Bame, Tony Beck, Woody Blanchard, Charles Fox, Peter Illick, Harley Michaelis, Forrest Miller, Ted Off, Mike Reed, Tyson Sawyer, Chuck Anderson, Norman Anderson, Jerry Arana, Bruce Baker, Ken Bates, David Batey Jr., Rich Border, John Boren, Mike Chiddick, Doug Dorton, Roger Egginton, Dale Folkening, Harlan Halsey, John Hohensee, Dave Jones, Stan Koch, Terry Luckenbach, Carl Mohs, Lee Mur­ray, Mark Nankivil, R. J. Ostrander, Jef Raskin, Les Rogers, Joe Ruminski, and Karl Widiner.

Prof. Mark Drela of M. I.T. is gratefully acknowledged for making his ISES computer code available to aid in the analysis of the new airfoil designs. Finally, MKS Instruments, Inc. and Scientific Solutions, Inc. are acknowledged for their valuable contributions of instrumentation.