chapter in the following figures:
5. DATA ANALYSIS Of the eight wind tunnel test runs performed, six trials had meaningful data. These data for these six wind tunnel tests was processed and they are displayed in this chapter in the following figures: Figure 5.1: Section Lift Characteristics for the 1/5 Scale Ultra-Light Wing Model Figure 5.2: Section Drag Characteristics for the 1/5 Scale Ultra-Light Wing Model Figure 5.3: Section Pitching Moment Characteristics for the 1/5 Scale Ultra-Light Wing Model Figure 5.4: Drag Polar Characteristics for the 1/5 Scale Ultra-Light wing Model The raw wind tunnel data is listed in Appendix A. The equations which relate percent of range and scale factor readings into actual lift, drag and pitching moment forces were obtained from an AE 245 laboratory exercise. These equations and along with lift, drag and pitching moment equations were written into a basic program to speed up the data analysis program. The final output of this program gives the tunnel speed, Reynold's number and the wing lift coefficient, drag coefficient and pitching moment coefficient. The output listing for runs 3-8 are in Appendix A.
The lift coefficient-angle of attack curve is seen in Figure 5.1. Data from trials number 6 and 7 were plotted.
Although these two trials were performed at 122 and 67 feet per second respectively, the data compares very well. The lift coefficients at higher angles of attack for the high speed case lies below those for the low speed case. This most likely indicates that wing section deformation at higher speeds lowers the wing's lift producing efficiency.
An unusual characteristic of this lift curve is that there appears to be two different and distinct lift curve slopes.
Between -4 and +2 degrees angle of attack the lift curve slope is roughly 7.6 per radian. Between +6 and 16 degrees angle of attack the lift curve slope drastically drops to 1.8 per radian. This indicates that this wing section does not generate much incremental lift coefficient at high angles of attack. Also evident is that lift coefficient is very sensitive to angle of attack change at small angles Of attack. Another interesting characteristic of this wing section is the high lift at zero angle of attack. The angle of zero lift is approximately -5 degrees. Obviously this wing section generates a relatively large margin of positive lift at small negative angles of attack.
The drag coefficient-angle of attack curve is seen in 4 -- Figure 5.2. Data for this plot was taken from test run #3.
Minimum drag for this wing section occurs between -4 and -2 degrees angle of attack. It should be clarified that this drag is for the entire model and support mount! No tare runs were performed due to time restrictions, since most of the data runs were taken at low speeds and since the model is relatively large this wont create a significant error.
The drag bucket in this curve also seems fairly symmetrical between -12 and +8 degrees angle of attack. One interesting characteristic of this curve is the intense amplification of drag at large angles of attack. The drag reading at 20 degrees is a factor of 24 times larger than the drag reading at -2 degrees. This "amplification factor" in ordinary wings is usually not as large. This is perhaps caused by the wing fabric pocketing at high angles of attack and further destroying the air flow. Another possible theory is derived from the fact that the wing frontal area to tunnel test section area ratio is small at large angles of attack.
The airflow is constrained to this area, and normal flow probably cannot be achieved, and the air pressure is probably increased, thus the drag is increased. A third L_ mmw possibility of excess drag at high angles of attack could be due to the model flutter at these angles. The model was seen to flutter at -12 degrees and above +8 degrees angle of attack. Drag is known to increase with flutter.
The pitching moment-angle o£ attack curve is seen in Figure 5.3. Data for this plot was taken from test rum #7.
It should be reminded that this pitching moment data is W about the main model support mount which is located at .18c of t_he wing. Pitching moment data is usually referenced at .25c or the aerodynamic center. A simple transformation can be performed to shift the pitching moment coefficient to this point but time constraints limited this process. Never the less, the slope and shape of the pitching moment curve is accurate and can be commented on. The slope of a pitching moment-angle of attack curve should ideally be a straight line. The pitching moment curve plotted indicates three different upwardly sloping "troughs". The angle of attack breaks between the three troughs are 0 degrees and 14 degrees. It is uncertain what causes these distinct breaks, but again it is assumed to be the fabric flexure.
_a Apparently fabric flexure change at 0 and 14 degrees angle of attack is very critical to pitching moment characteristics of the wing.
The lift coefficient-drag coefficient curve is seen in Figure 5.4. Data for the two curves were taken from test runs #6 and #8, the high speed and low speed trials, respectively. The slope of this curve indicates the maximum PW lift to drag ratio of the model. For the low speed case (run #8) the maximum lift to drag ratio is 12. The maximum lift to drag ratio for the high speed case (run #6) is 7.
This indicates that the lift to drag ratio is reduced at higher speeds. This is probably because the fabric flexure at higher speeds is more warped and less conducive to lift.
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IAE 592
ACTERISTICS FOR THE 1/5 SCALE ULTRA-LIGHT WING MODEL A_O APPO UNIVERSITY OF KANSAS _'AOE I,_ D w u m w CALC REVISED DATE FIGURE 5.3 SECTION PITCHING
AE 592
CMECK MOMENT CHARACTERISTICS FOR THE 1/5 SCALE _LTRA-LIGHT WING MODEL APPD APPD UNIVERSITY OF KANSAS '*_ i_ m m,,- L_ REVISED : DATE FIGURE 5.4 DRAG POLAR CHAR-
AE 592
ACTERISTICS FOR THE 1/5 SCALE CHECK I ULTRA-LIGHT WING MODEL APPD I APPD I PAGE I UNIVERSITY OF KANSAS _ 6. WING FABRIC FLEXURE The topic of wing fabric flexure was mentioned often in the previous chapter. The section shape of an ultra-light wing is highly variant to airspeed and angle of attack.
Airspeed tends to vary the magnitude of the fabric flexure.
Angle of attack varies the location and direction (inwards or outwards) of fabric flexure. The fabric flexure for five different angle of attack settings were sketched in Figures 6.1 to 6.5. The many different (and odd !) airfoil shapes should be noticed for the range of attack angle settings.
These figures show generalized airfoil shapes. The wing model was constructed with wire cross braces on the lower surface between the leading edge and main spar for fabric support (as stated in the construction chapter) which obviously are reflected in the lower surface fabric flexure shape. These helped to limit the fabric deflection in that particular area, but the exact shape they create is not determined in the figures.
-I0 degrees angle of attack: This setting is shown in Figure 6.1. The upper surface leading edge and trailing edge are indented signifying a pressure force exerted downward on the wing. The entire lower surface is bubbled outwards, again displaying a downwards pressure force.
There is a very interesting bubble in the fabric on the upper surface of the wing at about .25c. This perhaps is the only upwards pressure force on the wing, and serves to form a very unusual airfoil surface.
-6 degrees angle of attack: This setting is shown in Figure 6.2. The upper surface leading edge and trailing edge are indented, and so is the lower surface trailing edge. These indented surfaces are all handling inward pressure forces. The surfaces bubbling outward (experiencing outward pressure forces) lie on the middle upper surface and the lower leading surface of the wing.
0 degrees angle of attack: This setting is shown in Figure 6.3. The upper surface leading edge and entire lower surface of the wing are experiencing inward pressure forces. The remaining upper surface is bubbled outward and is experiencing lift.
6 degrees angle of attack: This setting is shown in Figure 6.4. It is virtually identical to the setting of zero degrees in Figure X.4. The only difference is that the upper surface fabric bubbling is more marked.
20 degrees angle of attack: This setting is shown in Figure 6.5. This is quite similar to the previous two settings (0 and 6 degrees), however the upper surface leading edge and lower surface fabric deflection is more marked, and the upper surface bubble is shifted more aft.
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w == fv i 7. RECOMMENDATIONS A_D COMCLUgZ0NS This project is an initial attempt to provide aerodynamic data for an ultra-light wing. Conclusive and fairly accurate lift, drag and pitching moment data were recorded and analyzed for the wing model. Some of the important findings are: I) The lift coefficient-angle of attack curve indicated the presence of two entirely different lift curve slopes at different angles of attack.
2) The change in drag between small and large angles of attack is quite marked.
3) There occur two distinct break points on the pitching moment coefficient-angle if attack curve, indicating a particular sensitivity at these two angles of attack.
4) Lift to drag ratios for this model are 12 at low speeds and 7 at high speeds.
5) Aerodynamic data for an ultra-light wing is a function of the fabric flexure, which in turn is directly related to angle of attack and airspeed.
There are range of other tests that could be performed with this wing model. Hopefully a structural failure test will not be one of them. Ideas for future experiments with this wing may include: I) Building a rigid model of the ultra-light wing to provide base data so that a more accurate study of the effects of fabric flexure can be studied.
2) Re-doing the drag data and taking drag tare data.
w 3) Calculating the pitching moment about a more useful reference point such as 0.25c.
4) Performing this testing in a different wind tunnel that can register the maximum forces endured by the wing.
Overall this was a very enjoyable project and it is encouraged that other students use this wing in individual or group testing--such as an AE 245 laboratory exercise.
Appendix A:
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REYNOLDS NUMBER = 333876 ALPHA I CL I CD CM 20 1.482431 .6364076 -.3382829 18 , 1.393485 .5403461 -.3179859 16 1.3'93485 .440282 -.2909233 .3582294 -.2638607 =7 14 1.334188 12 1.260066 .3061961 -.2537122 10 1.200769 .2561641 -.2455934 8 1.148884 .2261448 -.236798 6 1.074763 .1761128 -.2252'364 4 .9858166 .1360872 -.20499'34 2 .8523978 9.406024E-02 -.17929 w 0 .6300331 6.804358E-02 -.1420788 -2 .3483713 5.203333E-02 -.1204287 _'I II "_ .....
.i00_'_ -4 .1111823 3.202051E-02 -.1008083 -6 -.0667094 7.404743E-02 -7.645194E-02 -8 -.1111823 .1080692 -4.465334E-02 -it} -.3483713 .1561 -1.353132E-02 -12 -.4521415 .2181397 4.262365E-02 _'_ WIND TUNNEL RUN NUMBER 8 TUNNEL VELOCITY IN FT/S = 47.32559 REYNOLDS NUMBER = 237222. 1 ALPHA I CL I CD CM _ 20 1.5¢34973 .6640236 -.3350507 .545094 18 1. 394853 -.3048962 16 1.394853 .4757184 -.2948446 14 1.358147 .3429137 -.237886 12 1.32144 .2874132 -.2278345 I0 1.248027 .2457879 -.2211335 8 --1.233344 .19_8216 -.2144325 6 1.167272 .1585728 -.2010304 4 1.¢327787 .1129831 -.1789171 2 .8956426 7.730424E-02 -.1460821 0 .6240133 .049554 -.1172677 -2 .418456 .049554 -.0971647 -4 .1835333 4.360752E-02 -8.175237E-02 -6 -1.468267E-02 5.351832E-02 -.6499984 -8 -.1908747 8.325072E-02 -3.685558E-02 -10 -. 3670667 .1466798 -6.701014E-03 -12 -.4771867 .2180376 5.360811E-02 k 1 E V. STATIC TEST OF AN ULTRALIGHT AIRPLANE Reprinted from I. Aircraft, VoI. 25, No. 1, January 1988 Howard W. Smith Professor Department of Aerospace Engineering University of Kansas L_ i m - ! = = Partially supported by NASA Langley Research Center Grant #NAG 1-345 L_ H W VOL. 25, NO. 1, JANUARY 1988 J. AIRCRAFT 37
Static Test of an Ultralight Airplane
Howard W. Smith" University of Kansas, Lawrence, Kansas This paper describes the work mceesmu'y In perform the static test of an uitralight airplane. A steel reaction gantry, landing whiffletrees, hydraulk actnadoa system, and Instrumentation _slems were designed. Load and stress analyses were lamdormedon the airplane and on the newly designed gantry and whiffletrees. Load cell calibration and pressure Indicator calibration procedures are described. A description of the strain and deflection mea.,mrememsystem is Included. The engine, propeller, fuel, and pilot were removed and replaced with masses to fulfill center-of-gravity requlcements prior to testing. Data obtained to date are compared to the analytical predictions.
Nomendatur¢ Analysis CL = wing lift coefficient Design Criteria d ffi displacement, mm In the early days, an airplane had to be able to carry the F_ = ultimate compression stress, ksi Limit load without permanent deformation and the ultimate h = altitude, ft load for 3 s passing the static test sequence was a time of joy Mx = wing bending moment, N-m and celebration for the structures engineers. Nowadays, air- n = limit load factor craft are governed by much more rigorous specifications.
RN = nose wheel reaction, Ib The static strength requirement has been retained, but is now Rz = left main wheel reaction, lb only one element of a much larger array of specifications RR = right main wheel reaction, Ib under a comprehensive umbrella known as the structural S = wing area, ft 2 integrity program. Among the factors included are: corrosion, V = airplane speed, ft/s durability, damage tolerance, and flutter. Aircraft that are to Wo = empty weight, lb be certified prior to use must meet or exceed specifications.
WaF = basic flight design weight, Ib These requirements are specified in either Federal Aviation Regulations or Military Specifications and the "meet or exceed" phrase is satisfied by analysis or by test or both.
A set of design guidelines for an utralight has been published by the Powered Ultralight Manufacturers Associa- Introduction tion (PUMA). ( However, there are no specifications govern- ing the structural integrity of an ultralight airplane. For this S the service life of the fleet of ultralight vehicles increases, the number of fatal accidents is expected to analysis, the ultralight was treated as though it were a normal increase as well. Several cases have been documented by the category general aviation airplane governed by FAR-23. All National Transportation Safety Board1 in which the integrity related Mil-Spees and Mil-Standards were invoked as well.
of the structure was questioned. When similarities between It should be noted that student interest in this research cases occur, it is logical to formulate a plan to investigate the project was very high. One student elected to write a report on basic behavior of a typical vehicle.
a structural integrity program for uitraiights,S probably the The opportunity to formulate a plan presented itself in early only one of its kind in existence.
1985. Research on the aerodynamics and flight characteristics of an Airmass Sunburst "C" was drawing to a close and a master's thesis by Blacklock was published. Consequently, a full-scale ultralight airplane was available for further research.
A proposal was written and presented to the NASA-Langley Research Center. The primary goal of this proposal was to perform a structural test to destruction of an ultralight airplane.
The structural floor and the ultralight airplane specimen are shown in Fig. 1. To perform a static test, a steel gantry and its sway bracing was designed. 3 Similarly, the upper and lower •'hiffletrees were designed and integrated with the loading de- vice. Finally, the strain and deflection systems were designed.
This paper describes the details of the work accomplished.
Presented as Paper 86-2600 at the AIAA General Aviation Technology Meeting, Anaheim, CA, Sept. 29-Oct. !, 1986; received Oct. 28, 1986; revision received June 12, 1987. Copyright © American Institute of Aeronautics and Astronautics, Inc., 1986. All fights reserved.
"Professor, Aerospace Engineering. Associate Fellow AIAA.
Fig. ! Sunburst "C" ultralight.
H. W. SMITH VOL. 25, NO. 1 Table t Lift distribution Table 2 Weight breakdown of test ah'cruft, ib Structure Speed (maneuvering) 69 ft/s Altitude h 10130ft Tube WG-i 5.31 Wing skins 16.25 Weight WaF 468 Ib CL (max) 1.48 Landing gear S 150.9 ft a Wheel-nose 3.12 Main wheels and tires 10.90 n (limit) 3.8 Rear axle 7.01 Seat 8.71 Powerplant Lift Distribution Engine and propeller 78.38 Muffler 5.70 Ordinarily, a structural test engineer begins with air load Propeller shaft 8.88 distributions as "known" values. Both spanwise and chord- Misc., each < 3 lb Remainder wise pressure distributions must be given beforehand to allow We Weight empty 277.48 determination of "patch" loads. For this ultralight, six Fuel 15.52 spanwise and two chordwise stations were selected to simulate Pilot ("Bellerophon") 175.00 the subsonic pressure distribution. In reality, the airfoil W_ Basic flight weight 468.00 behavior is unknown, since it is only sail cloth stretched over the front and rear spar tubes. During a maximum positive load i factor condition, the airfoil is taut and has a particular set of ordinates. During any other flight condition, including inverted flight, the ordinates are variable.
\
Since an air load distribution was not available, one was 2.0 calculated using a quasivortex lattice method. This work was
\
done by a student who favored this method and the analysis was performed with ease. 6'7 With this knowledge, patch loads could be determined. Those data were incorporated in the 1.5 upper whiffletree design. The design maneuvering speed at a U, Limit load factor of 3.8 was 69.0 ft/s. (See Table 1.) The spanwise lift distribution is shown in Fig. 2. The spanwise drag distribution was assumed to be negligible.
1.0 Dead Weights
t The weight breakdown for our test condition is given in
Table 2. The engine, propeller, shaft, and mounts were ,__0.5 removed and replaced with a mass whose magitude and center of mass were correctly located. The lower whiffletree mass was included to correct the lg dead weight loads. Fuel was replaced with water of the correct weight.
0.0
t
Our ultralight pilot, named Bellerophon, was constructed o_- 0.0 0.2 0.4 0.6 0.8 1.0 army coveralls, worn-out army boots, a cap, and a mask" SPANWISE NON-DIMENSIONAL COORDINATE (Halloween) for cosmetic purposes. The cap was adorned with
Y
a NASA logc_ Bellerophon's center of gravity was built up with concrete cylinders at the buttock and thigh locations. The Fig. 2 Wing spanwbe lift eoeffldent.
remainder was constituted from plastic bags and Kaw River sand. Weighing and loading him into the aircraft required the assistance of four strong students. The upper whiffletree arrangement for the left-hand wing is Overall airplane weight and center-of-gravity location was shown in Fig. 4.
checked and rechecked by actual weighings with three balance The lower whiffletree is a loading mechanism as well. A scales under the wheels. Results of the weighings were: pair of steel straps connect at the engine mount holes and the RN--11.49 lb, RL = 127.0 lb, RR = 133.2 Ib, for a total of U-straps bear directly on the fuselage cage tubes. These 271.69 lb. (See Fig. 3.)
whiffletrees are commercial grade steel and are designated ders 6 and 7. Tier 6 is adjacent to the fuselage and tier 7 (the t Point Load Calculations lowest) connects to the 10,000 lb hydraulic actuator. A load cell is in series with the actuator. These linkages are bolted With many scientific developments, the creators of the directly to a floor fitting where they are reacted. The floor breakthrough cannot foresee the eventual applications of their fitting, called the "alligator," was specially designed for that work. Likewise, Joseph Fourier could not have known that his work with sines and cosines would be used to calculate air load purpose. It is located directly below the air load center-of- pressure vector P, shown in the lower whiffletree sketches pressures on an ultraiight airplane nor could Fred Whipple (Figs. 5 and 6). All of the lower whiffletree members are made have known that his method would be used to approximate from standard AISC steel sections: rectangular tubing, tees, that air load.
and flat straps.
The upper whiffletrees are simple three-point beam pairs made from_or_inary 2x4 and 2x6 pieces of lumber. There Internal. Loads Analysis are five ;'tiers7 of trees. The first is the highest and the fifth w the lowes_'.. The trees are connected with heavy-duty turnbuck- A stress analysis of the wing structure was performed using les. Tier 1 is connected to the steel gantry with a single steel the air loads discussed above. Availability of the Polo finite-element method and its ease of use were the reasons for strap. Tier 5 is just below the wing and is in direct contact with t the tubular spars. Plywood bearing plates are used to spread its selection, s Results are given in DeAlmeida's report, s The the load along the spars. Tiers 1-3 are the spanwise trees, flying wire loads at the design limit toad factor of n = 3.8 are:
!
while tiers 4 and 5 assure the chordwise center-of-pressure Forward inboard 44 lb location. With no load in the actuator, the uhralight is Aft inboard 65 Ib suspended above the hangar floor in straight and level flight.
JANUARY 1988 ULTRALIGHT AIRPLANE 39 Fill. 3 Weight and center of gravity.
z(up) TIER o TIER 7 ...
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Fig. 5 Lower whiffletree, left slde view.
Y c L AIRPLANE o X(FORWARD) -.v- _ ELASTIC AXIS UP
i
Fig. 4 Upper whiffletree.
L_
= OOI'BOAILD m Forward outboard 222 Ib Aft outboard 145 Ib Wing bending moments Mz and spar displacements d are shown in Figs. 7 and 8.
Systems Design For this study, the test rig was divided into four independent systems. The design and assembly of each system is described below.
Hydraulic System A 3000 psi hydraulic system was designed to apply the Ioad.
An Allis-Chalmers 10,000 lb, 8 in. stroke actuator and a Prince hand pump were purchased from a surplus machinery supplier. A pressure gage and short hydraulic lines were obtained from the same supplier. A schematic of the hydraulic Fig- 6 Lower whiffletree, rear view.
system is shown in Fig. 9.
The Boeing Company supplied the hydraulic lines, a eye shaft could then be gripped in test machine jaws. Excellent four-port BarksdaJe valve, and several hydraulic fittings. The linearity was achieved. A calibration constant was determined 2 gal reservoir and hydraulic oil were purchased locally. These to be 82 lb per unit readout. 9 parts were assembled and the lines purged of air by two students. The system was tested during the two-by-four destruction test described below. Deflection Measurement System Large deflections were measured with a sliding scale system.
Load Cell System In hazardous situations, a telescope or transit was used. This A 5000 lb Baldwin-Lima-Hamilton load cell has been in the was the case when cable failures were imminent. When Aero Department for a number of years. A pair of load cell deflections were small (less than l in.), a dial indicator was "eyes" had to be purchased to match the special internal used. Tip deflections of 3.70 in. limit were expected. The sliding scale concept was proved during the wood bending threads. The eyes have 1 in. diameter self-aligning bearings. A pair of links connect to a smaller eye at each end. The smaller destruction test, which was recorded on video tape.
P VOL. 25, NO. 1 H. W. SMITH m 4O 100 iiiiiiiiiiiiiiiiiilllll I 40 ll,lil i,, [,,,,ll,,,I,,,I I | 8o I,-,, t,-- Z 20 _ Z
° 7T
_ 6O / -2o- _'_// \\ II - - j e_ 40-- Front Spar // Z , Rear Spa e_ F =
7 Spar V -
-Io0 _,ii[,,,,l,t,,l,r,,l,,,, O.i 0.2 0.4 0.6 0.8 1.0 0.0 0.2 0.4 0.6 0.8 l.O SPANWISE NON-DIMENSIONAL COORDINATE SPANWISE NON-DIMENSIONAL COORDINATE Bg. 8 Wingllmitdefl_tlons.
Fig. 7 Wing limit bending moment*.
Strain Measurement, System ALLI S-CHAU4ERS ONE GALLON All strain gages were single-element foil gages from Micro RESERV.
Measurements. A 10 channel switch and balance unit and a _L GAGE, ICARSH -- X J-167B \ strain readout unit were available from previous research. The \ 3.000 PSI HOSE TENSION strain gage terminal board was borrowed from the Aerospace Medical Research Laboratory. The resulting strain measure-
0 12;,% I
ment system design was proved during the tube tension t ISUPPLY _ II i __'.'2.'ZZ-_ _.% .. II PN-HP-IO _ _ _tl_l_2._51Um --= component tests described below. Data were taken with a Vishay-EUis switch and balance unit and strain indicator.
Component Tests ' "' BARKSOALE LOW PRESSURE RETURN 14Z e3 AC3 Tube Compression Compression tests of the 0061-T6 tubes were run to verify Fig. 9 Hydraulic system.
the heat treat level. The ultimate stress in compression was: 2) Almost nothing is known about the behavior of an Fcu (measured) -- 47.8 ksi and Fcu (MIL-HDBK-5A)=42.0 ultralight structure under repeated loads. A durability and ksi.
damage tolerance research program is highly recommended.
Wood Beudlng Wood bending tests were performed on a pair of medium- , '!
Acknowledgments grade "S-P-F" lumber.The test simulated an upper whiffle- Many people freely volunteered to work on this project: tree and was performed tospot check themodulus of rupture Steve WaddeLl, Geoffrey Smith, Rou Schorr and Paul of "spruce-pine-fir," another unknown. Both the stress Oelschlaeger. Thanks to the Caroline Wire and Rope magnitudeand thefailure mode weremissed. The modulus of Company who supplied the cable and assembled the test rupture in bending, not to be confused with the civil specimens at no charge. This work was supported by NASA engineering designvalue, was estimated to be 9600 psi. The Langley Research Center under NASA Grant NAG 1-345 wood beam ensemble failed in horizontal shear and "prying" and the Aerospace Engineering Department of the University near thepoint of maximum moment. The magnitude was 85% of Kansas.
of the predicted ultimate load.For this test, the load-deflec- tion curve was linear up to 50% of thefailure load.
References Cable Tensiou _"Safety Study: UItralight Vehicle Accidents," National Trans- Cable testing was very interesting and informative. Four portation Safety Board, Rept. NTSBISS-85/01, Feb. 7, 1985.
2Blacldock, C. L. Jr., "Summary of the General Powerplant, assemblies of ¼ in. diameter, 7x 19 aircraft cable, were Weight and Balance and Aerodynamic Characteristi_ of an Ultralight designed to represent the "flying wires'on the ultralight. They i Aircraft," M.S. Thesis, University of Kansas, Lawrence, Aug. 1984.
were fitted with thimbles, grommets, tangs, and Nico-prefd _Smith, H.W., "Design of Static Reaction Gantry for artUltralight clamps. Failure load for the cable is estimated to b.e 1740 lb.
Airplane Destr_ction Test," AIAA Paper 85-4022, Oct. 14, 1985.
None of the cables carried more than 975 lb. All "failed" by *"Airworthine_ Standards for Powered Ultralight Vehicles,'" the cable sliding out of the Nico-press fitting. Cable testing i$ Powered Ultralight Manufacturers Association, Annandale, VA, Dec.
incomplete at this time. ALl cables will be fitted with double 9, 1983.
clamps and retested in an attempt to rupture the cable strands. vrurnipseed, Michael E., "Aircraft Structural Integrity Program for OltralighL_," University of Kansas, Lawrence, May 7, 1986.
Special safety precautions have been taken to keep humans _DeAlmeida, S.F.M., "Aerodynamic and Structural Analyses of an out of a 100 in. cable whipping lethal radius drawn with each Ultralight Aircraft," University of Kansas, Lawrence, May 6, 1986.
cable end as an arc center.
_Lan, C.T., "A Quasi Vortex Lattice Method in Thin Wing Theory," Journal of Aircraft, Vol. !1, 1974, p. 518.
tLopez, L.A. et al., "Polo-Finite," University of Illinois, Urbana, Recommendations 1985.
1) Unscathed portions of the ultralight, such as the wing *Page, L., "Cable Testing for Ultralight Airplanes," University of Kansas, Lawrence, May 6, 1986.
tip, can be sawn off and used in future wind-tunnel work. The t°(All) Engineering Drawings, University of Kansas, Lawrence, (125 two-dimensional lift and drag coefficients should be obtained drawings total).
from minimum to maximum CL.