In this module week, you will apply your gained knowledge about unaccelerated aircraft performance. Since you will build on your previously derived drag data for your example aircraft.
For your independent project, create an instructional presentation. Using/building on your previous drag (i.e., thrust-required) table and graph created in Module 3, generate additional power-required values in the table and depict the power-required curve for your aircraft.
Then, working with your derived thrust-required and/or power-required curves and table data, explain how to find various performance aspects for your aircraft, and provide the specific data for your example. At a minimum cover the following:
- Maximum range airspeed
- Maximum endurance airspeed
- Best climb conditions
- Best rate of climb (ROC) & associated airspeed
- Best angle of climb (AOC) & associated airspeed
- Maximum forward airspeed
- Best glide airspeed
Additionally, discuss and highlight numerically on a specific example of how weight change influences performance events such as the best range or endurance.
As in previous assignments, you will need to research additional information such as required formulas and pertinent aircraft data. Again, the emphasis in this project task is on explaining your methodology as if you are attempting to instruct someone unfamiliar with the aerodynamic details and relationships. Therefore, make sure to detail all assumptions, all formulas used, and all steps that were taken. The following will give you some starting points for your search and consideration.
- Required formula:
- Thrust to power relationship
- Weight change influence on performance airspeeds
- ROC & AOC relationships
- Necessary aircraft information:
- Powerplant output (for simplification, you can assume constant power output at the rated value across the entire speed range); whatever powerplant data you utilize, please make sure to include a short discussion detailing your assumptions.
- Previous information:
- Make sure to detail again all assumed information used/transferred from last week (e.g., aircraft weight, atmospheric conditions, etc.) since performance data are, obviously, only valid for specific cases and conditions.
As previously stated, you are encouraged to utilize appropriate computational software such as Excel® or MatLab®.
- Your presentation is due by the last day of the module and should be created using Powerpoint or the presentation platform of your choice.
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Independent Project: Drag
For the speeds in the first column, start with your aircraft’s stall speed, then continue in intervals of 20 or less knots (consider increasing the detail in the important portions near (L/D)max – see also this module’s tutorial videos within the 3.2 Lectures and Tutorials ), and continue to at least a speed of 300kts or higher if required to allow for answering the questions and explaining all drag phenomena.
To create this table you can use the Excel Spreadsheet ASCI 309 Table Drag Table (XLSX) located under “Field and Presentation Resources”.
|V (KTAS)||q (psf)||CL||CDP||CDI||CD||CL / CD||DP (lb)||DI (lb)||DT (lb)|
To fill out your table and subsequently create a diagram with the total drag curve, you will need to research a variety of variables, formulas, and components. Again, the emphasis in this project task is on explaining your methodology as if you attempted to instruct someone unfamiliar with the aerodynamic details and relationships. Therefore, make sure to detail all assumptions, all formulas used, and all steps that were taken. The following will give you some starting points for your search and consideration.
1. Assumptions and conditions:
· Assumed atmospheric conditions
· Calculated dynamic pressure (second column; based on the assumed atmospheric conditions and KTAS)
2. Necessary aircraft information:
· Wing size and configuration (e.g., AR & efficiency factor – if you can’t find an efficiency factor for your aircraft, you can make an assumption [i.e., pick a value] somewhere between 0.75 and 0.85)
· Weight (should, of course, fall between MTOW and empty weight of your aircraft)
· Airfoil information (e.g., CLmax from last module & CDP – if you can’t find the CDP for your entire aircraft, you can utilize the minimum drag for your airfoil and add a value of 0.02, which will account for the parasite drag of your aircraft’s fuselage) or if you are still having trouble, just use a Cdp of .021 which is a common Cdp.
3. Required formula (for inputs see the formula summary )
4. Resources and Inputs page)
· Dynamic pressure
· Lift equation (two forms: one solved for stall speed and the other solved for required CL)
· Drag coefficients (CDi & CD)
· Application of coefficients to find actual forces (Dp, Di, Dt)
· Possibly wing geometry conversions (e.g., wingspan and area into AR or wingspan and average chord into AR)
5. Do not forget to create the diagram.
Once created, utilize your derived table and diagram data to answer the following associated questions:
A. What are the minimum drag parameters for your aircraft?
· Minimum drag value D(min) 96.0 lbs
· Speed VD(min) at which minimum drag occurs 66.5 knots
· Relationship between Dp and Di at D(min)
parasite drag is about 5 lbs less than induced drag
B. What are the maximum lift to drag ratio CL/CD parameters for your aircraft?
· (CL/CD)max value 18.28
· Speed at which (CL/CD)max occurs 66.5 knots
C. Compare answers in A. and B. and comment on the findings.
(CL/CD)max takes place at similar airspeed of 66.5 knots at the occurrence of minimum drag
D. Explain which of your derived values will allow glide performance predictions for your aircraft and quantify best glide conditions with specific values.
When the aircraft is flying at 66.5 knots, the lift to drag ration is at its lowest. This enables the aircraft to produce the highest amount of lift through the use of the least amount to climb. This is effective for gliding conditions.
Airspeed Vs Drag
V(KTAS) 0 60 80 100 120 140 160 180 200 220 240 260 280 300 Dp(lb) 45.87 54.08 66.38 103.72 149.35 203.29 265.52 336.05 424.87 502 597.41999999999996 701.14 813.15 933.47 DI (lb) 50.2 43.8 34.700000000000003 22.2 15.4 11.3 8.6999999999999993 6.8 5.5 4.5999999999999996 3.9 3.3 2.8 2.5 DT (lb) 96 100 101 125.4 164.8 214.6 274.2 342.9 420.4 506.6 601.29999999999995 704.4 816 935.9
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