Further tunnel qualification tests with calibration models having similar frontal area to the 3.
The experimental uncertainty in the performance coeffi- cients was estimated using the methods of Kline and McClin- tock Ref. These uncertainties take in account uncertainty in the forces and moments determined from the balance calibra- tion and the uncertainty in the dynamic-pressure measurement. They do not include potential uncertainties in the model reference area, mean aerodynamic chord and wing span as these were assumed to be negligibly small.
All of these uncertainties were acceptable for the purposes of this investi- gation. In addition to this analysis, several repeat runs were performed for both clean and iced configurations and these run-to-run variations in the coefficients were much smaller than these uncertainties listed in Table II. The wings were first covered with self-adhesive Ultracote Plus surface covering.
This covering material provided an exceptionally smooth surface for the oil to flow. A base coat of SAE 30 weight motor oil was applied to the surface covering. Any excess was wiped clean prior to the application of a mixture of baby oil and UV fluorescent dye. The oil-dye mixture was applied with a small sponge roller used for trim painting applications. UV blacklights were used to illuminate the test surface.
The appropriate angle of attack and sideslip were set and the airflow was set to a dynamic pressure of 5 psf for approx- imately 10 min. The oil flow was monitored via video camera during the run and recorded on 8 mm tape. Thus, ambiguities were avoided by watching the development of the oil patterns during the runs.
Run times of 10 min were sufficiently long for steady state conditions to develop in the oil patterns. At the conclusion of the run, the fan was shut down and the tunnel was entered to take photographs with a digital camera. Usually the oil-dye mixture could be redistributed for the next run using the sponge roller.
After about four runs, it was necessary to wipe off the oil-dye mixture and apply a fresh coating. Artificial Ice Shapes It was determined at the beginning stages of this program that large, glaze-horn type ice accretion with the ice- protection system not activated should be used for aerody- namic testing on the 3. This was primarily due to the practical limitations of the 3. Because of the small model size relative to foil scale, smaller ice shapes would have been impractical to fabricate and install accurately. Also, previous research has shown that geometric scaling is appropriate for large, leading- edge ice accretion.
For example, Lee et al. They found little variation in the iced-model performance over this range with the large, leading-edge ice shape. Since time and resources did not allow for a suitable geometry and Reynolds number scaling study on a wing or tail surface component, a large ice shape was selected since it required neither.
The artificial ice shapes tested on the 3. This process was used to rapidly survey a large variety of potential icing-cloud conditions to produce a large glaze-horn ice shape. These were lofted into a quasi-three-dimensional geometry along the span of the wing and tail surfaces. Cross-sections of the ice shapes at various spanwise locations on the wing are shown in Figure 3. These cross-sections were taken in the streamwise direction and normalized by the corresponding local chord.
This corresponds to a length of 0. These cross sections show that the glaze ice shape becomes smaller at locations farther inboard. This is explained by the smaller total Figure 4. The ice shapes also tend to angle downward from the midspan station inboard because of the increased sectional lift values moving inboard on the wing. The ice-shape cross sections for the horizontal and vertical tail are shown in Figure 4 for the midspan stations of each surface.
For these figures the local chord length was also normalized by the wing mean aerodynamic chord length. The ice shape on the vertical tail was nearly symmetric since there was no net lift on this surface. The cross-sections shown in Figure 3 and Figure 4 are fairly smooth. This smoothing resulted from the lofting process that was used to develop the fully three-dimensional shapes from the LEWICE3D cross-sections. These three- dimensional shapes were manufactured using the fused deposition modeling technique and were fixed to the leading edges using a combination of alignment pins and tape as shown in Figure 5.
No roughness was applied to the smooth shapes. Note that there was no ice shape on the wing behind the engines. Results and Discussion Clean Model The static aerodynamic measurements were carried out for a large range of incidence and yaw angles to document the clean performance of the baseline 3. These data were compared to previous results for the 5.
The 5. In absolute terms, the lift-curve and pitching- moment slopes for the 5. This could be explained by the lack of blockage correction in the present data. However, the drag data would be similarly affected, while the data in Figure 6 match very well in the pre-stall region.
This agreement in drag also suggests minimal effect of the boundary-layer trips applied to the 5.
Figure 6. Stall for the 3. This is consistent with the drag coefficient that was lower for the 3. Of course, there was very little variation with angle of attack due to the model symmetry. The data show the large variation in these coefficients and change in sign with sideslip as expected for the lateral-directional Figure 8. The agreement in rolling-moment coefficient between the two data sets is fairly good over the range shown. It is unclear as to why there are larger differences for yawing moment.
In both cases the absolute slope is again lower for the 5. Similar comparisons were observed for the numerous incidence and sideslip angle sweeps performed. These plots are not reproduced here because similar results were observed for the present series of tests. At this angle of attack, the surface flow was fairly uniform in the streamwise direction everywhere on the wing, except near the trailing-edge tip region.
In this location, oil accumulated and appeared to flow upstream right at the trailing edge of the wing tip; thus indicating local boundary-layer separation as indicated in the figure. This complex flowfield is also described with the aid of Figure 10 which is a schematic representation adapted from Poll Ref.
This vortex drew free-stream fluid down to the surface along the reattachment line indicated in Figure 9 and Figure This formed a closed, leading-edge separation bubble with surface flow moving upstream between the reattachment location and the leading edge. This is indicated in Figure 10 and can be seen in Figure 9 as the line of oil accumulation upstream of and approximately parallel to the reattachment line.
Adapted from Poll Ref. Thus indicating a slight decrease in the size of the leading-edge separation bubble. The flow visualiza- tion in Figure 1 1 indicates that the spanwise vortex had broken down in to two separate structures. There appears to be a small spanwise-running vortex from the fuselage -junction region forming a partial-span separation bubble. Then on the outboard side of the pylon another partial-span vortex appeared to have formed. Unlike at lower angles of attack, the vortex in this case appears to have lifted off the surface in a downstream direction near the wing midspan section.
From the midspan section inboard to the root, the surface flow is generally in the streamwise direction to the trailing edge, indicating significant loading on this inboard half of the wing. This is seen most clearly on the outer 25 percent of the span where the oil flowed upstream from the trailing edge of the wing. Still visible near the leading edge is the secondary separation of the reverse flow on the outboard half of the wing. Poll Ref. This had the effect of increasing slightly the portions of the wing with reverse surface flow. All of the images shown in Figure 8, Figure 9 and Figure 10 were for the right wing.
Flow visualization was also performed simulta- neously on the left wing. No significant differences were observed. The iced aerodynamics of the GTM was investigated with the artificial ice shapes installed on the wing and tail leading edges. The incremental increase in drag due to the ice shapes is clearly seen over the angle of attack range in the figure. In contrast, the lift data showed virtually no effect due to the ice shapes. In fact, there may have been a slight increase in maximum lift and stalling angle in the iced configuration.
Little difference was also observed in the pitching moment. While it appears that the effect of the artificial ice shapes on the GTM performance was minimal, this appearance was likely due to Reynolds number effects on the clean model. It is well known that the performance of two-dimensional airfoils and three-dimensional wings is influenced by Reynolds and Mach number Refs.
For Mach numbers less than about 0. As discussed later in this sections, Reynolds number effects on iced-airfoil and wing performance are much smaller in contrast to the clean configuration. Kaneshige Ref.
These data are also plotted in Figure When comparing the estimated full-scale clean GTM data with the subscale iced-GTM data, the expected decease in maximum lift and stalling angle for the iced configuration is clearly seen. This could have been due to the contribution of induced drag resulting from the higher lift coefficients for the clean configuration.
The effect of the artificial ice shapes on the lateral-directional characteristics of the 3. At this sideslip angle, Shah et al. The results are further mixed for the sideslip angle sweeps for the two incidence angles shown in Figure However, the yawing moment variation was mostly unaffected by the artificial ice shapes.
P deg. Figure In contrast to the clean, or un-iced, configuration, Reynolds number effects on lift and drag coefficient for airfoils, wing and airplane models with artificial ice shapes have been shown to be small. Aerodynamic testing was conducted with three different model scales: full scale, 42 percent scale and 8. Three different ice configurations were investigated, including a large glaze-horn similar to that used in the present work.
Lee et al. In contrast, there was a large change in performance of the clean wing with Reynolds number as expected. These data are reproduced here in Figure 15 and Figure The clean wing data in Figure 15 show the classic effects in the significant increase in maxi- mum lift coefficient and stalling angle with increasing Reynolds number.
For Reynolds numbers greater than 4. Typical Reynolds number effects were also seen in the pitching moment variation with angle of attack that grew more linear in the pre- stall region with increasing Reynolds number.
For example, the dragonfly, Sympetrum sanguineum [ 54 ], has an aspect ratio of 9. There's a steering control in the cockpit, but that's the only thing a plane has in common with a car. Imagine two air molecules arriving at the front of the wing and separating, so one shoots up over the top and the other whistles straight under the bottom. White, F. Deep learning. Why is there a downwash rather than simply a horizontal "backwash"?
This behavior is contrasted against that shown in Figure 16 for the same wing model with a large glaze-horn artificial ice shape similar to that used for the present tests. There was considerably less dependence of the performance data on Reynolds number for the iced-wing configuration. Therefore, it is reasonable to conclude that the lift and drag performance of the iced 3.
The percent decrease in lift coefficient and percent increase in drag coefficient due to the artificial ice shapes was com- puted using the present data for the iced-GTM configuration and estimated full-scale clean performance shown in Figure These increments are plotted versus angle of attack in Figure This behavior is typical of icing flight encounters where the primary effect is reduced airspeed in cruise or holding conditions. The reduction in maximum lift coefficient and stalling angle of course become more important as more lift is required for lower airspeed conditions.
Also shown in Figure 17 are analogous data for the semi-span business jet wing tested by Lee et al. In both cases, a large glaze-horn type ice shape was used in the iced-wing configuration. The data for the semi-span wing model show similar percent increases in drag, but much larger reductions in lift coefficient. Figure 1 7. Such factors notwithstanding, however, further comparison does yield some perspective on the present data. For example, Reehorst et al. Two full-span, leading-edge ice-shape configura- tions were tested, although detailed geometry information was not provided.
For one configuration that appeared to be similar to the iced-GTM configuration in the present tests, the reduction in C L at stall was only 12 percent based upon the low-Reynolds number data. This is low relative to the values plotted in Figure 17 and was most likely due to Reynolds number effects for the clean airplane model.
Fligher Reynolds number, iced-configuration data for the twin-engine, commercial transport category airplanes similar to the present GTM and that of Reehorst et al. Lynch and Khoda- doust Ref. However, Zierten and Hill Ref. For the ground frost simulations, the entire upper surfaces of the main wings were covered with appro- priately sized roughness on the flight test airplanes and wind- tunnel models. Van Hengst et al.
There is a wide range of maximum lift reductions from 10 to 35 percent, depending upon the airplane configuration and whether the data were from flight or wind- tunnel test. For the Boeing models, the flight test points and wind-tunnel data produced similar maximum lift reductions despite large differences in roughness size. Based upon these data, the 24 percent lift reduction in the present tests on the GTM cf. Figure 17 may be on the low side, given the large size of the artificial ice shape relative to the roughness sizes shown in Figure These roughness configurations may be useful for benchmarking the iced-GTM performance in any follow-on testing, provided suitable simulation methods can be implemented.
An additional important distinction between the present data and that shown in Figure 18 for full-scale airplane configurations is the presence and use of the high-lift system. Such systems were not incorporated into the 3. These flow visualization images were compared to those from the baseline configuration to note any significant differences due to the artificial ice shapes. Figure 8. Similar surface-flow topology was observed comparing Figure 19 and Figure 9 for the clean model. The exception was that the approximate bubble reattachment location laid much closer to the leading edge in Figure 19 than in Figure 9.
Figure 20 shows that there appeared to be two distinct flow regions, one inboard of the engine pylon and another outboard. The partial-span vortex on the inboard portion of the wing is not as visible in Figure 20 as in Figure 11, but the flow was generally in the streamwise direction toward the trailing edge. Larson, L. The Airplane Experiment. Physics Teacher. Churchill ; Illustrated By Jim Michaels.
The use of BLC in modifying aerodynamic characteristics. Weltner, K. Utilizing simple and inexpensive equipment, elementary and middle school science teachers can conduct. This teaching guide includes a brief history of rockets,. The fan will be your power source, but you need a method for creating a laminar flow field.
The easiest way to do this as an amateur is to use a grid of stacked straws, like a honey comb, inside its own frame. Basically, you build a square frame and apply some insect screen to one face. Assemble the straws in the frame, then cover the opposite side of the frame with insect screen. Place the fan against this frame.
Now build a test area inside a long square tube that attaches to the other side of your flow straightening section the frame with the straws is the flow straightener. This reduces the swirl caused by the fan. There are some rough rules for "correct operation. Minimize interior projections that means try to make the inside as smooth as possible. This reduces entrained turbulence the rules are actually alot more specific, but for a school project, this should get you started.
Look up "flow straightener", "laminar flow", and "wind tunnel" good luck. Tip: To turn text into a link, highlight the text, then click on a page or file from the list above. This Sidebar appears everywhere on your workspace. Add to it whatever you like -- a navigation section, a link to your favorite web sites, or anything else. Aeronautics Page history last edited by Joel Rosenberg 6 days, 21 hours ago.
LAB: A student guide for aeronautics. Publisher: New York, Hastings House, inc. Haaren High School. Authors: Housel, David C. Pittsburgh, Pa. The Paper Airplane Book. Numerical Calculation of Model Rocket Trajectories. Jenkins, R. Measuring Model Rocket Acceleration. Gerhab, G. Boundry Layer Control on Airfoils. Brna, P. Strickler, Jr.