# aerodynamics homework help

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• Q1: =) Spanwise pressure coefficients are shown for 3 streamwise stations of a delta wing at two angle of attack conditions, a = 6.8° (top row) and a =16.3° (bottomrow). From these plots, explain the vortex behavior for each condition. See Answer
• Q2:As "task1.m" script runs from "P2amain.m" script, it must output the following two plots (no user inputs required): (i) The "stagnation streamline" plot (the shape of the semi-Rankine body: r/R) in the range of: 09⁰, 179° (upper half surface) and = 181°, 191°, 201°,... 351° (lower half surface). 19°, 29°,. (ii) The "pressure coefficient" plot (the surface pressure distribution over the semi-Rankine body: C₂) in the range of: 0=9°, 19°, 29°,... 179° (upper half surface) and = 181°, 191°, 201°,... 351° (lower half surface). (Note) the C, plot in aerodynamics is "negative values in positive y coordinate direction" (rule of thumb). This will require you to modify the y axis data direction ("flip-flop" of the y coordinate) when you perform C, plot in MATLAB.See Answer
• Q3:You have a CubeSat of 10×10×30 cm³ in size and 4 kg in mass. Estimate the size of a launch vehicle that is capable of inserting this CubeSat (4 kg payload) into a circular LEO of 200 km-altitude. The primary question is how small it could be when it is built using available technology today.< Conditions: 1) You are free to use any tools available for your design and/or analysis. Even a hand-calculation is fine.< You are free to choose any type of propellant available. You may use known values of any performance parameter without a detail calculation as long as you declare the source.< Consideration of launch site, launch angle, number of staging, drag loss, gravitational loss, and so on is up to your choice. Just make it clear what kind of assumption is made in your estimation.< You may choose any launch site with any launch angle, however, it might be suggested to select either of launch site of a) Naro Center in Southern Korean Peninsular or b) US Cape Canaveral, Florida with launch angle 180 degree or to the straight south direction.<See Answer
• Q4:Part 1: Assume the A320neo is currently flying with an airspeed (TAS) of 250 kn at an altitude of 5000 ft and is climbing. The throttle factor is set at F, = 0.85. The mass of the aircraft is 78.5 tonnes. Determine: a) the climb angle and rate of climb. b) the steepest climb angle and maximum rate of climb. c) compare the current airspeed of 250 kn with those required to achieve the steepest climb angle and maximum rate of climb.See Answer
• Q5:Part 2: The aircraft reaches an altitude of 30000 ft and begins cruising. The mass of the aircraft at the start of cruise is 78 tonnes and the fuel load that can be used for cruise is 12 tonnes. a) assuming a cruise at constant altitude, determine the cruise range as well as the optimal airspeed and Mach number at both beginning and end of cruise. b) assuming cruise at a constant airspeed (given by the optimal airspeed at the beginning of cruise) with a drag ratio n = 3, determine the cruise range as well as the altitude the aircraft reaches at the end of cruise. c) assuming cruise at a constant airspeed (given by the optimal airspeed at the beginning of cruise), what would be the tailwind speed required to increase the range to 4700 km (assume tailwind remains constant throughout the cruise)?See Answer
• Q6:t Figure 8.4 shows a drag polar plot of a glider, with C₁/CD and C2/C3 superimposed as func- ons of the lift coefficient. For a given weight, W, wing area, S and altitude, h (this determines the density), by selecting a range of values for angle of attack, it is possible to compute: C₁/CD, C/C, V and RD. The corre- sponding performance data are shown in Table 8.1. The glider selected for this example is a relatively poor one by modern standards. Its lowest flight path angles are high compared to those achievable with modern gliders. Trimmed lift-to-drag ratios of 40.0 to 50.0 have already been achieved, yielding descent angles as low as 1.2 degrees! Despite this, the reader is asked to verify that the approximations of Eqns (8.22) - (8.24) are quite good! The performance results of Table 8.1 are plotted in Figure 8.5. The reader is asked to check out the significance of the points labeled A, B and C in Figures 8.4 and 8.5. These very important points are discussed next.See Answer
• Q7:The reader is asked to show, that the drag polar of Figure 8.4 can be approximated by a parabolic drag polar with A=15.6, e=0.82 and C= 0.0150. The speed for best glide performance, using sea-level density data can then be computed from Egn (8,29) as 20.9 m/sec. This agrees closely with the data of Table 8.1 which use the actual drag polar.See Answer
• Q8:This is the first part of a project in Aerodynamics. Project is about creating a flying device which uses aerodynamics of a maple seed You have to find information about maple seeds You have to submit a one page document in APA format which describes: (Problem statement) which includes (Inspiration, design, use, how) Some information regarding the project topic: https://www.imnovation-hub.com/science-and-technology/biomimetic- drones/#:~:text=A%20drone%20inspired%20by%20the,range%20of%20a%20lightweight%20drone https://m.youtube.com/watch?v=Rm_KoTBtVz4&pp=ygUSbWFwbGUgc2VIZHMgZmx5aW5nSee Answer
• Q9: Question 1. Calculate the temperature (in Kelvin) at an altitude of 9665 metres above sea level, assuming ISA atmosphere: O -47.82 O 0.0065 O 350.82 O 12 O 225.18 Not answered Question 2. Calculate the temperature (in Kelvin) at an altitude of 11471 metres above sea level, assuming ISA atmosphere: 0.0065 216.50 O 3 O 362.56 O-59.56 Not answered SESSMEN sity The Questions Question 3. Calculate the pressure (in Pascals) at an altitude of 5192 metres above sea level, assuming ISA atmosphere: O -19 O 52611 O 90423 O 53 101325 [a2-alkarbi] UFMFRU-15-1 DEWIS E-ASSE Not answered The Questions Question 4. Calculate the air density (in kg per cubic metre) at an altitude of 12504 metres above sea level, assuming ISA atmosphere: O 81.64 O 0.29 O 0.36 O 288.00 O-80.99 [a2-alk UFMFRU-15-1 DEWIS E-AS Not answered The Questions Question 5. An aircraft is flying at an altitude of 8364 m above sea level. Its airspeed with respect to the surrounding air is 137 m/s. Assuming ISA conditions, calculate the dynamic pressure (in Pascals). O 1965.1 O 288 -39.37 4716.52 O-928.7 OFMPRU-15-1 DEWIS E-ASSESSMEN Not answeredSee Answer
• Q10:Evaluation and Discussion 8. Task A: A1: A2: A3: A4: A5: A4: A5: Why pressure distribution on the upper and lower surface are the same for NACA0015 airfoil at zero AOA? Why there is a difference in pressure distribution between the upper and lower surface for cases you simulated in this Task? Which sides (upper or lower surface) has higher pressure, and why? Do you see the difference in pressure (between upper and lower surface) changes with AOA? Why? Describe the pressure distribution on the upper surface by identifying the stagnation point, suction peak, and adverse pressure region. What is Cp? How do we compute the lift coefficient of the aerofoil from the Cp distribution? (15 marks)See Answer
• Q11:9. Task B: B1: B2: B3: 15 marks) B4: As small AOA, how do you see C₂ varies with a ? As AOA increases, does the relationship you observes in B1 continues? Why? Is there any difference in the C₁ vs a relation you observed in B1, for airfoil with difference thickness, and different camber? At even higher AOA, what happen to the relationship between C₂ and a? In your own words, describes the flow phenomenon that has occurs. B5: Tabulate and discuss the Stall AOA and maximum C₂ of the 4 airfoil studied. B6: Compute and discuss the lift-curve slopes for the 4 airfoil studied. (15 marks)See Answer
• Q12:10. Task C: C1: C2: C3: Based on your simulation result, identify regions where boundary layer is laminar, turbulent and transition. How do you identify transition point and from you simulation result discuss how they are affected by AOA and Re. How to you identify separation point and discuss how they are affected by AOA and Re. (20 marks)See Answer
• Q13:11. Task D: D1: D2: D3: D4: Why the sectional lift coefficient C, is zero at the wing tip? Why the sectional lift coefficient C₁ is almost "flat" near the mid-section of the wing? What is the cause of C, distribution? Do you think the C₂ distribution is elliptical? If not suggest ways to make it elliptical. Use a few data point to verify the induced drag formula. (20 marks)See Answer
• Q14:12. Reflection: In less than 100 words, write a short reflection and summary on what you have accomplished and learnt in the Laboratory assignment. (20 marks)See Answer
• Q15:Question 1 (10 marks) (2022 Paper) You are working as a flight test engineer at SUSS Research Flight Test Centre. During the low speed phase of the test program for a UAV program, pressure at two locations were measured. The first pressure is known to be at the stagnation point, and the second is at the edge of the boundary layer above the wing. You know that the UAV is flying at an altitude of 500 m under the ISA condition. The two pressure readings are: Point 1: Stagnation Pressure, pi=96,920.125 N/m² Point 2: Static Pressure just outside boundary layer, p2=94,454.204 N/m² Question la Apply the flight test data provided to calculate the velocity of the UAV, and determine the Mach number and the Reynolds Number of the UAV (assume that the characteristic length is 0.2 m and the Dynamic Viscosity is 1.7737 X 10³ kg/ms). Question lb Verify the compressibility of the flow over the UAV, and state whether the density can be treated as a constant. Question le Examine if Bernoulli's equation could be applied to determine the local flow velocity at point 2, and state the reason(s). Question Id Appraise the flow velocity at the point 2. Question le Appraise the flow around the point 2 to determine the surface pressure on the wing, directly below point 2.See Answer
• Q16:Question 2 (10 marks) (2022 Paper) Question 2a The Aerodynamic characteristic of a NACA0015 airfoil is simulated using a Potential Method. The Cp and Tw distributions are presented in Figure Q2(a)-1 and Q2(a)-2 of Appendix 1 respectively. Question 2a(i) Verify boundary layer transition occurrence by identifying the location of the transition Point. Question 2a(ii) Base on the data presented, provide TWO (2) reasons to verify your answer in Q2a(i). Question 2a(iii) Apply the learnt airfoil content, is NACA0015 a symmetrical airfoil? Question 2a(iv) Appraise the Lift-curve Slope (Cle) of this airfoil. Question 2b The airfoil presented in Q2(a) above is used for the design of a wing with Aspect Ratio of 10. Apply the data to calculate the Lift-curve Slope (Cla) of the wing. You may assume the span efficiency factor to be 1.0.See Answer
• Q17:Question 3 (10 marks) (2022 Paper) Question 3a The drag polar of a propeller powered aircraft is given as: C₂ = 0.0095 +0.038 x C² The aircraft weights 10,000 N and has a wing area of 8 m². Question 3a(i) Analyse the speed of the aircraft at sea level when Power required is minimum. Question 3a(ii) If the engine power at sea level ISA condition is given as 80 kW. Due to the air traffic congestion near the aero dome, the aircraft was asked to climb at its maximum climb rate. Set up the appropriate method to calculate the following: (1) (2) (3) The climb speed. The maximum Rate of Climb. The corresponding climb angle. Question 3b During the test flight, the following data were measured by the test boom installed at the nose of the aircraft: OAT (outside air temperature) = 240 K P = 37,759 N/m² Po = 44,790 N/m² Question 3b(i) Apply the given data to calculate the Mach number of the aircraft. Question 3b(ii) Appraise the True Air Speed (TAS). EAS301 Copyright © 2022 Singapore University of Social Sciences (SUSS) TOA- July Semester 2022 Page 5 of 15/nQuestion 3b(iii) Appraise the Calibrate Air Speed (CAS) the pilot might read from the cockpit Air Speed Indicator (ASI). Question 3b(iv) Appraise the Dynamic Pressure at the condition of the flight.See Answer
• Q19:5. Format of the report The report has a strict 4-page limit. Anything above four pages will be penalised according to the marking guidance. The report should contain a front sheet (which will not count towards the page limit), containing the following information only: Name Student ID Module Code Module Title • Coursework Name The report should be written as a formal lab report with appropriate headings. Do not include a table of contents or appendix. Do not repeat large sections from this handout in relation to the background or the procedures. As a rough guide, approximately three pages should cover the raw and processed data, with one page to cover a discussion of the flow physics. Note that the raw data should be presented to you in the units it was. Appropriate references are essential to show that you have engaged with and understood the data and the theory. The required format is single-spaced Times New Roman font, size 12 pt. The page margins should be set to normal, i.e. 2.54 cm at Top/Bottom/Left/Right.See Answer
• Q20:Question 1 2 pts Question 1: In a wind tunnel test of flow over a wing section, air enters the test section at a speed of 54 m s1 and a pressure of 942 kPa. Assuming Bernoulli's principle applies to the flow, what is the pressure at the stagnation point on the wing surface if the air density is 1.16 kg m 3? The answer should have units of kPa.See Answer

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