FORMAT OF THE PROJECT REPORT
DOCUMENT # 1
EXTERNAL AND INTERNAL LOAD CALCULATION FOR THE F-16 WING (Reports are due April 3rd Wednesday. Presentation day will be announced later)
DATE SUBMITTED: TITLE OF THE DOCUMENT: Give an appropriate title GROUP MEMBERS:
INTRODUCTION Give basic information about the F-16 fighter aircraft. This information should contain performance/mission characteristics of the aircraft, structural layout of the wing/fuselage etc. You are supposed to search internet for this. Describe the tasks accomplished in the first document such as how you performed the external and internal load calculation. Summarize the output you produced.
Given information:
Necessary information for the load calculation of the F-16 wing is given below.
Wing Area = 27.88 m2 Aspect ratio = 3.2 Taper Ratio = 0.2 Leading Edge Sweep Angle = 40 deg Root Airfoil = NACA 64A204 Tip Airfoil = NACA 64A204
Flight Design Gross Weight (of the whole airplane) = 26910 lb Limit load factor: 9 Ultimate Load Factor = 13.5 (9x1.5)
Airfoil data: (Check if the airfoil data given below is okey or not. Search in the internet for NACA 64A204)
NACA Section 64A204 (Stations and ordinates given in per cent of airfoil chord) Upper Surface Lower Surface Station Ordinate Station Ordinate 0 0 0 0 0.5 0.4052 0.5 -0.2717 0.75 0.4925 0.75 -0.3183 1.25 0.635 1.25 -0.3792 2.5 0.9042 2.5 -0.4714 5 1.3029 5 -0.5731 7.5 1.6156 7.5 -0.6358 10 1.8798 10 -0.6817 15 2.3062 15 -0.7407 20 2.6363 20 -0.7732 25 2.89 25 -0.7854 30 3.0785 30 -0.7804 35 3.2074 35 -0.7592 40 3.2775 40 -0.7198 45 3.2794 45 -0.6513 50 3.218 50 -0.558 55 3.1015 55 -0.4489 60 2.9353 60 -0.3314 65 2.7223 65 -0.2117 70 2.463 70 -0.096 75 2.1579 75 0.0046 80 1.8018 80 0.0702 85 1.3753 85 0.0701 90 0.931 90 0.0517 95 0.4673 95 0.0241 100 0 100 0 L.E. radius = 0.123 percent c slope of mean line at LE = 0.0951
1- CAD MODEL OF THE OUTER MOLDLINE OF THE F-16 WING
Based on the information supplied create CAD model of the outer moldline of the F-16 wing. You can draw only one wing. Make sure that reference area of the CAD model (half wing area) matches the information given above. If necessary you can change the sweep angle of the leading edge to match the 27.88 m2 total wing area.
You are free to create the CAD model of the outer moldline in any CAD software such as CATIA or even in Patran.
2- CALCULATION OF EXTERNAL AERODYNAMIC LOAD
External aerodynamic loading and its spanwise distribution will be calculated by using two different cases assuming a positive symmetric maneuver at load factors of 9 and 13.5.
Case 1: In this case, assume that aircraft is flying at subsonic speeds and exhibiting a symmetric pull-up maneuver at a load factor of 9.
For this case, the spanwise variation of the wing loading should be calculated by the ESDU document 95010 ‘Computer program for estimation of spanwise loading of wings with camber and twist in subsonic attached flow’ No twist should be taken into consideration. However, camber effect should be included in your load calculation. Note that ESDU calculates the lift and moment with respect to quarter chord.
This document should end with the plots of spanwise lift and pitching moment distribution calculated at the Mach number 0.8 and assuming sea level flight.
* Determine the angle of attack to fly at 0.8.
* Carry out spanwise lift and pitching moment distribution at Mach number 0.8 by the ESDU code 95010. Give the input/output of ESDU code. In this step calculate the spanwise variation of the loading (lift and pitching moment) first separately for the angle of attack and camber, and finally for the overall angle of attack and camber effect.
Make tables and graphes for the overall spanwise lift (local over all lift coefficient) and pitching moment coefficient distribution (local overall pitching moment coefficient). Show how you incorporated the camber effect.
Note: In the input file of the ESDU code, If Lm is selected as 1 then loading is due to incidence only. If Lm is selected as 3 loading is due to camber only. If Lm is selected as 1 and 3 then loading is due to incidence and camber both. In addition, P specifies the calculation mode. You can specify it as 1 so total loading calculation is also made.
For the load calculation due to camber only, set NL=1, Lm=3 and P=0. For the load calculation due to incidence only, set NL=1, Lm=1 and P=1 and specify your AOA or AOAs.
Case 2: In case 2, we assume that aircraft is flying at supersonic speed and making a symmetric pull-up maneuver at a load factor of 9 and 13.5. For supersonic speeds you can assume that pressure center is at 50% of the MAC.
Calculate the external aerodynamic load using Schrenk’s approximation for the symmetric maneuver positive load factor of 9 and 13.5 separately. Gross weight of the aircraft is 26910 lb (note this is the total weight of the airplane).
This method is commonly used to determine overall span-wise lift distribution, especially at the preliminary design stage.
The method states that the resultant load distribution is an arithmetic mean of:
• A load distribution representing the actual planform shape • An elliptical distribution of the same span and area
For the Schrenk distribution, draw only the spanwise lift distribution assuming that center of pressure lies in 50% chordline.
Comparison:
Compare your spanwise lift distributions determined by ESDU and by Schrenk’s approximation. Make two comparisons.
1- Compare spanwise lift distribution calculated by the ESDU code due to incidence only (do not include the camber effect) and by the Schrenk method for the 9g case. 2- Compare spanwise lift distribution calculated by the ESDU code due to incidence and camber and by the Schrenk method for the 9g case.
3- CALCULATION OF INTERNAL LOADS
Method 1: For the ESDU external loads, obtain the internal loads at Mach 0.8 and for the 9g case. Internal loads should be given as the spanwise distributions of shear force, pitching moment and bending moment, and they should be calculated in a similar fashion as the example we did in class. Also note that Bruhn’s example is also linked in the web site, so you can also follow that. For the number of stations, you can use the spanwise stations you used in the ESDU code. Make tables of the internal loads. Note that for the ESDU case, you obtain internal loads with respect to local quarter chord about which the overall lift and pitching moment coefficients are specified.
* Draw spanwise variation of the sectional shear force, pitching moment and bending moment curves. Comment on your results
* Check the accuracy of you results: Root shear force should be equal to half of your aircraft weight !!!
Note: During the calculation of the internal loads make sure that the running load is acting through the %25 chord along with the pitching moment. You must not carry the running load to any other chordwise position as the example by Bruhn shows.
Method 2:
For the Schrenk distribution, assuming that at the supersonic speeds, center of pressure is at 50% MAC, calculated the spanwise shear force and bending moment distribution at 50 % MAC. Do this in two ways:
1- Calculate the shear and bending moment distribution in the same manner as we have done in class using strips and strip loads for the 9g and 13.5g cases separately. 2- For the trapezoidal wing of F-16, integrate the Schrenk distribution to calculate the spanwise shear force and bending moment variation for the 9g and 13.5g cases separately.
* Draw spanwise variation of the sectional shear force and bending moment curves . for the 9g and 13.5g cases separately. Comment on your results and compare 9g results with the ESDU results for the 9g case.
* Check the accuracy of you results: Root shear force should be equal to half of your aircraft weight !!!
DISCUSSION AND CONCLUSION
Make any critical comments you have, discuss your results and conclude by summarizing the tasks accomplished.
PRESENTATION Groups should present first part of project in a single presentation. Presentation day will be fixed in class.