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Q1:4. Download the Matlab code posted to the Canvas (in supplementary materials) for integrating the problem of a compressible laminar boundary layer over a flat plate with zero streamwise gradients in the overriding freestream and with the viscosity following a temperature power law with exponent n = 0.7. The gas is calorically perfect and assume that the edge conditions involve a Mach number Me = 5. Use and modify this code accordingly to answer the following questions: a) Determine the ratio of the adiabatic wall temperature Ta,w to the edge temperature Te, along with the recovery factor r. State whether the approximation r = √Pr is a good one for these conditions. (10 points) b) For an isothermal wall at Tw/Te = 3, plot the dimensionless profiles of the static temperature T/T, stagnation temperature To/Te, density p/p and streamwise velocity u/ve, all as a function of the normalized wall-normal coordinate y √Reex/x, where x is the streamwise coordinate and Ree,x is the edge Reynolds number based on x. (30 points) c) For the conditions described in (b), calculate the products Cr√ Reex and CH√ Reex, along with the Reynolds analogy factor 2CH/C, where Cy is the skin friction coefficient and Cy is the Stanton number. Compare the Reynolds analogy factor with Pr-2/3. (10 points) Attach a screenshot of your modified version of the "main.m" script with your solutions.See Answer
Q2:3. The Blasius solution to the steady-state two-dimensional incompressible laminar boundary layer equation for uniform flow over a flat surface can be shown to be as follows: 4(x,y) = √vUxf(n) Where x and y are streamwise and wall-normal coordinates, respectively, v is the kinematic viscosity (which is a constant for an incompressible flow), U., is the freestream velocity, 4' is the stream function (which is related to velocity as u = 4/ǝy and v = -ǝч/ǝx in an incompressible flow), u and v are the component of the velocity in x and y directions, respectively, n = is the similarity variable, and f can be found by solving the following y √vx/U∞ ordinary differential equation (ODE): + Boundary conditions: f(0) = f'(0) = 0 and f'(∞) → 1 The solution to this ODE has the following asymptotic forms for small and large values of the similarity variable: {f(n) = 1² f(n) =n-√3 6 for n << 1 for >> 1 a) Using the above results, obtain an expression for the nondimensional surface shear stress, i.e., Tw, in terms of the Reynolds number based on the distance x. (5 points) b) Find the nondimensional vertical component of velocity at y→ ∞, i.e., " boundary layer in terms of the Reynolds number based on the distance x. (5 points) う in theSee Answer
Q3:1. A wedge with semi-angle = 20° is moving through air (T = 200 K, Pr = 0.7225 and y = 1.4) at a Mach number more than one. The wedge is kept at temperature Twall = 1600 K. Calculate the free-stream Mach number, Ma.., at which the wedge wall temperature becomes equal to the adiabatic wall temperature of the boundary layer. What is the corresponding Mach number at the edge of the boundary layer, Maę? Hint: Assume that the recovery factor is equal to √Pr, the gas is calorically perfect, and ignore the effect of boundary layer on the inviscid flow. (30 points) Ma∞ > 1 Tw B 20 shock 20B boundary-layer edgé u(y) Mae, Pes Tes HeSee Answer
Q4: SUSS SINGAPORE UNIVERSITY OF SOCIAL SCIENCES EAS309 Flight Dynamics and Control Lab Assignment Jan Semester, 2024 INSTRUCTIONS TO STUDENTS: 1. This is an Individual Assignment 2. This Assignment carries 15% of the marks toward the total Grade 3. You should prepare and submit this assignment as a formal report. 4. There are three submissions for this assignment : Submission Submission 1 Submission 2 Final 5. Submission Details: Weightage 10% 15% 75% Due Date 3rd March 2024 24th March 2024 6th April 2024 a. Submission via Canvas b. your submission should be either in .doc or .pdf format c. accompanying excel file or Matlab file to show your calculation could also be submitted. Introduction In this assignment, you are going to apply all you have learnt in EAS309 to analyse and evaluate the flight dynamic characteristics of a modern fighter aircraft. You will also attempt to design a simple flight control system to improve the aircraft flight characteristics so that it meets the flying quality requirement of a Specification (MIL- 8785). The F15 is a twin engine, highly manoeuvrable, supersonic, multi-role fighter aircraft. As the aircraft is highly manoeuvrable, the operating flight envelopes covers a wide range of AOA and AOSS. At these extremely high aircraft attitude, the aerodynamic flow over the aircraft is complex and has significant effect on the aircraft flight dynamic and stability. The variation of the key aerodynamic derivative coefficients are show in Appendix I. SUSS Aerospace System is requested to investigate the F15 flight dynamic characteristics at various AOA at both subsonic and supersonic speeds. Each student in the EAS 309 class will be assigned for one particular case, to study into the flight dynamic characteristics. You will received your assigned data package via email. The key parameters of the F15 is as follow: Mass S с b 20411.66 kg 56.485 m² 4.86 m 13.045 m Ixx Iyy Izz Ixz 38911.98 kg-m² 223845.45 kg-m² 254758.19 kg-m² -705.03 kg-m² Assignment Instruction Each student in the EAS309 will be receiving the F15 aerodynamic derivative data package. You may want to know that each data package is corresponding to different flight condition, i.e. for different Mach number and altitude, thus the data each student received will be different from each other. Based on the data package received, you shall perform the analysis, conduct the evaluation, and perform the flight control design as required in the questions listed in this assignment. You are required to work through this assignment using MATLAB. You may notice that the entire assignment span throughout the entire Semester, and you are strongly encouraged to do this assignment progressively, to avoid the last minute rush, pacing with the progress of the lectures. A recommended progressive schedule is shown below: Lab Lab #1 Lab #2 Lab #3 Lab #4 Lab #5 Lab #6 Dates 26 Jan 2024 16 Feb 2024 23 Feb 2024 8 Mar 2024 23 Mar 2024 6 Apr 2024 Activities Familiarization with MATLAB Read in parameters, static stability margin Compute all derivatives Submission #1 (3rd Mar 2024) (covering questions : 1A and 2B) Compute A and B, eigenvalue, eigenvector and modes of motion Compute aircraft response with control input Submission #2 (24 Mar 2024) (covering questions: 2C, 3B, 3C, and 4C) Simple feedback control State-space feedback, pole placement Finally, you are required to submit an individual Formal Report documenting all your findings and answers. The marks for respective and final submission are as follow : Submission #1: Only Questions underlined (10 Marks) Submission #2: Only Questions Bolded (20 Marks) Final Report: All Questions. (70 Marks) Assignment and Questions 1) Analysis the aircraft longitudinal static stability by answering the followings: (10 marks) A. Apply the appropriate theory to evaluate and present the numerical value of stability margin for your flight condition. B. From your result appraise the stability of F15, and state whether the aircraft is statically stable? 2) Appraise and evaluate all the stability derivatives (both longitudinal and lateral- direction axes) of the aircraft, and assemble the A and B matrixes (for both longitudinal and lateral-direction axes) of the aircraft (10 marks) A. Apply the theory learnt to describe the longitudinal and lateral-direction axes B. Analyse and Tabulate your answers for ALL the longitudinal and lateral- directional derivatives. C. Present your answer for Matrixes A and B for both longitudinal and Lateral-direction Equations. D. Appraise and explain the physical significance of matrix A and B? 3) Formulate and develop procedure to calculate the eigenvalue and the eigenvectors longitudinal and lateral-direction modes of the aircraft. for the (20 marks) A. Appraise and explain the physical significances of the eigenvalue and eigenvector? B. Analyse and present your answer for the eigenvalues and eigenvectors for both longitudinal and later-directional axes. C. Present the time response plot for all the aircraft Longitudinal and Lateral- Directional modes of motions D. Explain how you calculate the modes of the aircraft in Matlab. E. Apply the theory learnt to comment the lateral directional modes of the aircraft. F. Apply the theory learnt to comment on the longitudinal modes of the aircraft. G. If the aircraft was stable longitudinally, appraise the longitudinal stability characteristics : i. In terms of eigenvalues and eigenvectors ii. What are the two typical physical modes, and iii. What are the typical behaviour of these two modes. 4) Formulate and develop procedure to compute and plot the aircraft lateral- directional response when the aircraft is subjected to an impulse aileron input (10 marks) A. Apply theory learnt to explain the difference between impulse and step inputs? B. Appraise and explain the formulation you used in Matlab to compute the aircraft responses. C. Analyse and present the time response of the aircraft when subject to impulse aileron inputs D. Apprise and comment on these time response curves. i. Do you see an isolated mode or all the modes ii. Do you see any difference "when time is small" and "when time is long" 5) Formulate and develop a simple pitch feedback classical control is to improve the damping ratio of Phugoid mode. (20 marks) A. Appraise and explain the choice you have for the transfer function in your Matlab formulation B. Analyse and present the Root-Locus Plot C. Apply theory learnt, comment on how the damping ratio is changed with different gain, using the Root-locus plot computed. D. Appraise the meaning of "gain" in the question above.See Answer
Q5:2. Add an afterburner to the above situation. The maximum stagnation temperature downstream of the afterburner is 2000 K. The afterburner combustion efficiency nab is 0.95 and the total pressure ratio is rab is 0.97. All other efficiencies remain the same. Use a gamma value of 1.4 up to the primary burner, and a value of 1.3 for the rest of the engine. Assume R is 0.287 kJ/kgK throughout the engine. Plot the specific thrust, TSFC, nth, np and no as a function of rc, the total pressure ratio across the compressor. Consider a range of rc from 2 to 60. Assume the exhaust is ideally expanded. Also plot the nozzle area ratio as a function of rc. Comment on the changes that occur due to the addition of the afterburner (i.e. compare the plots to that of problem 1). Note that the overall fuel-to-air ratio can not exceed the stoichiometric value (f+fab<fst).See Answer
Q6: AERO 480 Assignment Problem Consider a jet trainer with following characteristics: Weight & Dimensions • Max weight - 7600 kg Wing Span -9.72 m • Wing Area - 23.52 m² Power plant • 2 Honeywell F124-GA-200 (twin engine) Maximum Thrust - 2* 2850 kg Aerodynamic CD 0.25 0.2 0.15 0.1 0.05 хо Y 0.003 0 -2 -1.5 -0.5 0 C -M-0.3 0.5 1 1.5 2 Using the data given above Plot the following: CL vs CD for M= 0.5 & 0.7 using the data given in the graph (Use compressibility correction) Thrust required at (a) sea level; (b) 10000 ft. • Power required at (a) sea level; (b) 10000 ft. • Excess power at (a) sea level; (b) 10000 ft. • . Excess power at (a) sea level; (b) 10000 ft. (if one engine failed & rudder has sufficient yawing control) Plot CL | CD' C1/2 C³/2 !²/CD' 12/CD vs Speed at (a) sea level; (b) 20000 ft. Calculate: • The Maximum speed at (a) sea level; (b) 10000 ft. • • The Maximum speed at (a) sea level; (b) 10000 ft. (if one engine failed & rudder has sufficient yawing control) The Stall Velocity (a) sea level; (b) 10000 ft. if CL,max = 2.0 The Maximum rate of climb at (a) sea level; (b) 10000 ft. The Maximum rate of climb at (a) sea level; (b) 10000 ft. (if one engine failed & rudder has sufficient yawing control) The Speeds at maximum rate of climb at (a) sea level; (b) 10000 ft. • • Service Ceiling of the aircraft Service Ceiling of the aircraft (if one engine failed & rudder has sufficient yawing control) • The minimum sink rate at (a) sea level; (b) 10000 ft. • The Speeds at minimum sink rate at (a) sea level; (b) 10000 ft.See Answer
Q7: CIVL 321 LABORATORY Spring 2023 Objective: Fluid Mechanics Lab #11 Aerodynamic Drag and Lift To determine the drag and lift coefficients of a model using the closed-loop wind tunnel with 28"x20" test section in LANG 122. Specifically, determine the drag and lift coefficients of a NACA 4412 airfoil- shaped wing model at relatively high Reynolds number(s) as a function of the angle of attack (a) and compare with published data. Theory: The total drag force of an immersed body consists of viscous drag (skin friction) and pressure (form) drag. Lift is solely due to pressure effects. The drag and lift coefficients and Reynolds number for airfoil- shaped wings are normally defined as: CD = 2FD PAU² CL = 2F PAU² Uc Re= V where FD is the measured drag force, FL the measured lift force, p is the fluid density, U the upstream fluid velocity (relative to a stationary body), A, is the planform area of the body (i.e., the area seen from above at zero angle-of-attack), c is the chord length of the airfoil-shaped wing, and v is the fluid kinematic viscosity. Refer to Sections 9.3 and 9.4 of your textbook for more information. Procedure: 1. If necessary, install airfoil mount and secure airfoil-shaped wing with screws with the help of the instructor. 2. Turn power to Force Balance Pitch/Yaw Control system ON. Release locking bolts on the wind tunnel force balance (do NOT lean on the frame) and then tare lift and drag displays. 3. The RP240 PITCH and YAW control panel should now be displaying a short prompting message. Re- tare the lift and drag displays if necessary. 4. Start the wind tunnel by performing steps 1 through 8 of the closed loop wind tunnel start-up procedure shown below. 5. Set the speed control to a value indicated by the instructor and record 1) the pitot-static tube deflection in inches of H2O; 2) the local static pressure in inches of H₂O; 3) the barometric pressure in kPa, and 4) the temperature in °C. The difference between the local static pressure and barometric pressure is a measure of how much the static pressure in the wind tunnel is below atmospheric pressure. This will allow you to calculate the air density in the wind tunnel using the Ideal Gas Law. Watch your units!!! 6. Select the PITCH axis (Press F1 on the RP240) and enter -12° (angle of attack). Record the drag and lift forces (lbs) indicated on the meters. 7. Increase the airfoil/wing angle of attack in increments of 2° up to +20°, recording lift and drag values for each angle. Return the airfoil to 0° angle of attack when finished. 8. To obtain lift and drag for other velocities, repeat steps 5 through 7 and record the necessary data. 9. Return airfoil to 0° angle of attack and follow the shutdown procedure for the wind tunnel below. 10. Press EXIT (F6) on the RP240 PITCH and YAW control panel. 11. Carefully secure all four locking bolts. Pay close attention to the load cell displays - try to keep displayed loads <2 lbs. as the bolts are being tightened. 12. Turn power to force balance control system OFF. Operating Procedure for Closed-Loop Wind Tunnel: Start-up: 1. Make a visual inspection to ensure no loose objects were inadvertently left in the wind tunnel and latch the door(s) shut. Scanned with CamScanner 2. Turn on the Test Section Lights. 3. Turn on the wind tunnel Master Power switch. The Wind Tunnel Speed Control should illuminate and control should initialize to the stopped condition. 4. Open (if not already) the Magnetic Clutch Water Valve (ball valve on left) and press the Clutch Cooling On button. Make sure water is circulating by checking the drain valve in the Southeast corner of the lab. 5. SLOWLY open the heat exchanger water inlet valve (it is the ball valve on the right) 6. Press the Motor Start button. The clutch output will slowly reach an idle speed of 250-300 RPM 7. Press P1 on the Wind Tunnel Speed Control to access manual control on the Run potentiometer. 8. Set the Run to desired RPM by adjusting the potentiometer and press Run. This initiates live control of the fan motor. For the first test use 800 RPM. When ready for the second test, SLOWLY adjust the potentiometer to 1100 RPM. Shut-down: 1. Press the Stop button on the Wind Tunnel Speed Control. 2. Allow the motor to coast to idle speed of 250-300 RPM. 3. Press the Motor Stop button. 4. SLOWLY close the heat exchanger water inlet valve (located on right) 5. When motor has coasted to a stop, push the Clutch Cooling Off button to stop the motor/clutch cooling water. 6. If the Magnetic Clutch Cooling Valve (ball valve on the left) was open when you started, then leave it open. Otherwise close it. 7. Turn off Master Power switch. 8. Turn off the Test Section Lights. 9. Turn off Main Electrical Cutoff Box. Velocity Measurement with the Pitot Tube: The Pitot tube, when properly connected to a manometer or pressure transducer, measures the difference between the stagnation pressure and static pressure of a fluid flow (AP). The velocity of the fluid at the static pressure tap, which is essentially the freestream velocity (U); can be determined from the Bernoulli equation: U = 2AP P = 2pmgh P where p is the density of the flowing fluid, pm is the density of the manometer liquid at standard conditions, and h is the deflection of the manometer liquid level read from the inclined scale. This scale is calibrated to read inches of H2O, so pm is the density of water at room temperature. The measured drag force includes both the drag on the wing and the vertical mounting strut. The drag on the mount was determined independently as: Fam = 1.4604×10-5V2.281, where V is measured in ft/s. Subtracting Fam from the measured drag force will correct for the additional drag on the strut so you obtain just the drag force on the wing. Material to be Included in Report in Addition to Normal Requirements: 1. Prepare an Excel graph of the lift coefficient (CL) versus angle of attack (a). Indicate the angle of attack at the point of stall (acmax) and the zero-lift angle of attack (ac=0). Compare with published results and discuss any differences. Compare the plots for the data collected at 70 mph versus 100 mph (you can plot both data on the same graph). Why does this result occur even though they are collected at different speeds? 2. Prepare an Excel graph of the drag coefficient (CD) versus lift coefficient (CL) and label each data point with the corresponding angle of attack. This is called a drag polar diagram (Figure 9.35b). Find the angle of attack where lift- to-drag ratio is maximum. Compare with published results given in your lab manual and discuss any differences. Note: Drag polar plots in many aerodynamics books are plotted with the drag coefficient on the y-axis and the lift coefficient on the x-axis. 2 Scanned with CamScanner Airfoil Designation: NACA 4412 chord: 5 in. span: 16 in. 800RPM 1100RPM Settings Test 1 Test 2 Yaw: 0 0 degrees Initial Pitch: 0 0 degrees Pitch increment: 2 2 degrees Pitot-static Pressure: 2,67 4.61 in. H₂O Local Static Pressure: 387 .710 in. H₂O Barometric Pressure: 10016 100.6 kPa Air Temperature:| 22.2 22.9 °C Tunnel Static Pressure: units Air Density units Air Speed units units Kinematic viscosity (air)| Reynolds No. (chord length): Test 1 Test 2 Angle of Attack, a Drag Force (degrees) (lbs) Lift Force (lbs) Drag Force Lift Force (lbs) (lbs) -12 1.212 2. h 2.242 4.47 -10 1.033 -1.33 1.874 -3.19 -8 0.831 1.20 1.438 -1.79 -6 2.723 10.48 1.298 -0.19 -4 0.677 0.32 1.262 1.50 -2 0.665 1.17 1.277 3.88 0 9.653 1.97 1. 330 5.68 2 0.708 3.10 1.447 7.45 4 0.776 4.01 1.597 9.13 6 0.854 4,84 1.776 10.58 8 0.972 5.61 1.998 11.97 10 1.101 6.26 2.235 13.12 12 1.224 6.79 2.532 14.99 14 1.350 7.10 2.353 14.73 16 1.541 7.39 4.210 11.09 18 1.729 7.43 4.592 11.10 20 2.338 5.37 4.803 11.17 Scanned with CamScannerSee Answer
Q8:3) Calculate the free rudder neutral point for the aircraft whose characteristics are given below. Determine its longitudinal static stability. dɛ = 0.4, 0.4, n=1.0 спос = -0.005 [1/º] (CL)=0.09 [1/º], (c) = 0.08 [1/º], =-0.003 [1/9], da Спо = 0.0, Cha x. ac = 0.5 0.8 Xcg x 0.3, S₁ = 3.5 m² S₁ = 10 m² 0.55 0 τ 0.35 0.7 Control Surface Area Lift Surface AreaSee Answer
Q9:1) Given a rectangular wing of aspect ratio 6 and area of 600 ft². The wing section employed is an NACA-4412 airfoil with an aerodynamic center at 0.24c and Cm ac= -0.088. If the wing is balanced so that the c.g. lies on the wing chord but 6 in. ahead of the a.c., calculate the lift/lift coefficient for which the wing would be in equilibrium (Cmcg=0). Is this lift coefficient useful? Is the equilibrium statically stable? Calculate the position of the c.g. for equilibrium at CL=0.4. Is this equilibrium statically stable?See Answer
Q10: ARO 3111 Computer Assignment Part I Conical Flow: Supersonic Flow Past Un-Yawed Cone Objective: Develop a computer program in a high-level computer language (MATLab, C++, etc.) to numerically calculate the properties of the conical flow generated by a right circular cone flying at a supersonic Mach number at zero pitch and zero yaw through an inviscid perfect gas. The cone has a sharp tip (vertex). Numerical Approach: To accomplish this, numerically integrate the Taylor-Maccoll (T-M) equation, which is a second order and very nonlinear ODE. Do so by first converting this second order equation into two first-order ODEs. Solve these two by the use of the powerful and popular algorithm: 4th-order (Classical) Runge-Kutta method. Use the "inverse" approach described in Anderson (Modern Compressible Flow, 2021, Section 10.4, pp. 373-374) by starting at the given conical shock wave angle 6, and marching in the negative -direction, in small increments of A from the shock towards the cone. O is the spherical angular variable, measured from the cone axis up to an arbitrary ray emanating from the cone vertex. Recall that "conical flow" is just the flow between the shock and the cone body, is fully isentropic, and is characterized by the remarkable characteristic that all flow properties are constant along each ray, though they vary from ray to ray. Of course, the free stream flow is isentropic ahead of the shock. The only region of non-isentropic flow is across the conical shock wave, which like all other shocks is adiabatic. Thus, (using Anderson's notation) T = constant everywhere, but po will drop across the shock, and will remain constant behind the shock. Also, note that, although, the shock wave is 3-D, the shock is locally planar, and thus locally can be treated by the use of 2-D oblique shock theory. Governing Equations: In Anderson (2021), T-M equation is Eq. 10.13, but instead work with the more convenient non-dimensional version of it in Eq. 10.15. You will also be using Eq. 10.14 to find the 0-component of velocity V. The latter will look the same in non-dimensional form, except for adding a prime on each side. Once you have found the two components of velocity, V, and V, you can combine them to find the magnitude of the local velocity vector V. From the latter, you can now find the local Mach number M from Eq. 10.16, which you will need for isentropic flow calculations in the conical flow field. For convenience, drop the primes from these three equations, i.e., Eqs. 10.14, 10.15 and 10.16. Notice a small misprint in Anderson p. 374, Fig 10.4: V-component of velocity just downstream of the shock should be parallel to and just downstream of the shock, not as shown. Notation: Use that of Anderson (2021): • Station 1 or ∞ is the free stream 8 • Station 2 is right behind the shock • Station c is at the surface of the cone • Subscript o is stagnation values • Cone angle is • Conical shock angle is ẞ or • Initial turning of the streamline as the flow crosses the shock from 1 to 2 is 8 which is <0 Case Study: 1. Run your program for: M₁ = 6 y=1.4 2 0. = 12° A0=-0.1° (this is small enough to give you good accuracy, we are marching in the negative direction, thus a negative sign on the step size). Find that produces this shock wave and the corresponding M. 2. Once your program is working, we will calculate and extract more data from it: calculate non-dimensional flow properties on each ray and plot as a function of Programming: See the attachment for the governing equations and the 4th-order Runge-Kutta (R-K) algorithm. In your computer program, use all the same notations as in the attachment. Structure your computer program with separate modules and each one should start with one or more comment statement(s) that describe what is being done in that module and also define your notation for the module. Note that you will need the free stream velocity nondimensionalized by Vmax, which is obtained from Eq. 10.16, applied to the free stream. You will see that the answer is not 1, but less than 1. You will need this value to compute your initial conditions at Station 2 to start your R-K computations. Hint: The final answers for 0, Cpc, and M can be read from Charts 5, 6, and 7 in NACA 1135, good checks on the veracity of the core of your program. 3 ARO 3111 Computer Assignment Part II Conical Flow: Supersonic Flow Past Un-Yawed Cone Now that your computer program is working, we will add some gas dynamics to it and extract some of the flow properties in the conical flow region, between the shock and the cone. Again, our Case Study is: M₁ = 6 y=1.4 0. = 12° A0=-0.1° You already found the values of 0 and M. for this flow from your two plots. Now, continue your work as follows: 1. Use locally planar oblique shock relations to obtain from flow properties at Station 1 (the free stream), flow properties at Station 2 (right behind the shock), as you did in Problem 10.1 of Anderson (2021) in PS9. Do your work in non-dimensional form, by hand calculation, with 4 significant figures, as well as by incorporating into your program, to find: M2, P2 P1, P2 P₁, T₁/T₁. They should agree completely. This also helps debug the early part of your program. 2. Now, recalling that flow from Station 2 to Station c is isentropic, use your previously developed program to compute and plot, in one color graph, local values of these properties as a function of the spherical variable 0, i.e., M, p/p₁, p/p₁, T/T. Theta will now go from the above to O, that is, an inverted scale. Add the results in (1) to complete your plot for theta between and 45°, to fully see what happens to the fluid particle as it approaches the cone. 3. Add one table of all three plots vs 0 in increments of one degree. 4. You will present your work in a report.See Answer
Q11: Lab 5 Advanced Plotting T5.1-1 Plot the equation y = 0.4V/1.8x for 0≤x≤ 35 and 0 ≤ y ≤ 3.5. T5.1-2 Use the fplot command to investigate the function tan(cos x) - sin(tan x) for 0≤x≤ 2. How many values of x are needed to obtain the same plot using the plot command? (Answer: 292 values.) T5.1-3 Plot the imaginary part versus the real part of the function (0.2 + 0.8i)" for 0 ≤ n ≤ 20. Choose enough points to obtain a smooth curve. Label each axis and put a title on the plot. Use the axis command to change the tick-label spacing. T5.2-1 Pick a suitable spacing for t and u, and use the subplot command to plot the function ze 0.5t cos(2016) for 0≤1≤8 and the function u = 6 log10 (²+20) for -8≤v≤8. Label each axis. T5.2-2 Plot the following two data sets on the same plot. For each set, x = 0, 1, 2, 3, 4, 5. Use a different data marker for each set. Connect the markers for the rst set with solid lines. Connect the markers for the second set with dashed lines. Use a legend, and label the plot axes appropriately. The rst set is y=11, 13, 8, 7, 5, 9. The second set is y =2, 4, 5, 3, 2, 4. T5.2-3 Plot y = cosh x and y = 0.5e* on the same plot for 0≤x≤2. Use differ- ent line types and a legend to distinguish the curves. Label the plot axes appropriately. T5.2-4 Plot y = sinh x and y = 0.5e on the same plot for 0≤x≤ 2. Use a solid line type for each, the gtext command to label the sinh x curve, and the text command to label the 0.5e* curve. Label the plot axes appropriately. T5.2-5 Use the hold command and the plot command twice to plot y = sin x and y = x-x³/3 on the same plot for 0≤x≤ 1. Use a solid line type for each, and use the gtext command to label each curve. Label the plot axes appropriately. T5.2-6 Plot the following functions using axes that will produce a straight-line plot. The power function is y=2x-05, and the exponential function is y = 10¹ x T5.2-7 Plot the function y = 8x³ for-1 ≤x≤ 1 with a tick spacing of 0.25 on the x axis and 2 on the y axis. T5.2-8 The spiral of Archimedes is described by the polar coordinates (0,r). where r = a0. Obtain a polar plot of this spiral for 0 ≤ 0 ≤ 4T, with the parameter a = 2.See Answer
Q12: كلية الهندسة College of Engineering جامعـة قـطـر QATAR UNIVERSITY 1. Objective: ● ● ● 2. Background: MECH 416 - Aircraft Design Assignment No (01): Aircraft Performance Analysis 3. Task: Understanding fundamental performance parameters Applying equations to calculate performance characteristics Interpreting and analysing performance results. Aircraft performance analysis is a critical component of aircraft design. It involves the evaluation of various factors such as aerodynamics, propulsion, and structures to determine how well an aircraft will perform under different conditions. The goal of performance analysis is to ensure that the aircraft meets its intended design requirements and is safe, efficient, and cost-effective to operate. It also allows for the optimization of the aircraft's performance, which can lead to improved fuel efficiency, increased range, and reduced maintenance costs. Overall, performance analysis is an essential part of aircraft design, as it ensures that the aircraft meets the needs of its intended users and performs reliably and efficiently. ● The data for the small single-seat home-built jet airplane are as follows: Wing span: 5.2 m. Wing planform area: 3.5 m². Gross weight at take-off: 4270 N. Fuel capacity: 55gal. (1 gallon of JP-4 ≈ 3.08 kg) Power plant: one French-built Microturbo TRS 18 turbojet engine with maximum thrust at sea level of 898.5 N and a specific fuel consumption of 1.3 N/(N.hr) the drag polar for this airplane can be estimated to be: ● In this assignment, you will make a Performance analysis for small single-seat home-built jet airplane, using the following information and requirements: ● جامعة قطر QATAR UNIVERSITY Cp = 0.02 +0.062C² Page 1 of 3 كلية الهندسة College of Engineering جامعة قطر QATAR UNIVERSITY a. the maximum velocity at sea level and b. the maximum velocity at 3.048 km. 1. Plot the thrust required and thrust available curves at sea level, and from these curves obtain the maximum velocity at sea level. 2. Plot the thrust required and thrust available curves at 3.048 km, and from these curves obtain the maximum velocity at 3.048 km. 3. calculate analytical (directly) Vmax 1/2 √(TA) [(TA)max/w] (W/s) + (W/s) √ [(T₁)max/w] -4CD,K Poo CDo Max Thrust available at altitude can be estimated by using this Equation (T^) max = (T₁). [206 ● Compare the result in 3 with those from 1 & 2. 4. calculate 2 6. Use the analytical results to calculate directly جامعة قطر a. The maximum value of CL/ CD b. The maximum value of C₁¹/2/CD c. The maximum value of C₁³/2/CD d. The velocities at which they occur at sea level e. The velocities at which they occur at 10,000 ft 5. Plot the power required and power available curves at sea level. From these curves, estimate the maximum rate of climb at sea level. a. Maximum rate of climb at sea level and the velocity at which it occurs. Compare with your graphical result from 5. b. Maximum climb angle at sea level and the velocity at which it occurs. 7. Consider our Aircraft flying at 10,000 ft. Assume a sudden and total loss of engine thrust. Calculate QATAR UNIVERSITY a. the minimum glide path angle, b. the maximum range covered over the ground during the glide, and C. the corresponding equilibrium glide velocities at 10,000 ft and at sea level. Page 2 of 3 كلية الهندسة College of Engineering جامعة قطر QATAR UNIVERSITY 10. Using the analytical approach described in Lecture Note 6 page 12, calculate the minimum time to climb to 10,000 ft. 8. Plot the maximum rate of climb versus altitude. From this graph, estimate the service ceiling. 9. analytically calculate the service ceiling, and compare this result with the graphical solution obtained in 8. 11. Estimate the maximum range at an altitude of 10,000 ft. Also, calculate the flight velocity required to obtain this range. 12. Estimate the maximum endurance. 4. Requirements: ● 5. Grading Rubric: جامعة قطر ● QATAR UNIVERSITY ● Choose either MATLAB or an Excel spreadsheet for your analysis. Use scientific principles and relevant equations to perform the calculations. Document your code with clear comments explaining each step and the formulas used. ● Accuracy: Correctness of calculations. Analysis: Ability to interpret results and draw meaningful implications (e.g., how do the results compare to typical aircraft of that class, or inform aircraft design choices). Clarity: Presentation of calculations and explanations. Page 3 of 3See Answer
Q13:/n/n Lab report format : Abstract Literature review Introduction Calculatoin Discussion References For graphs : Prefer matlab/excel Word limit: 1500 to 1800 words Referncing style : APA 7th editionSee Answer
Q14:Part 2: Written Project This assessment involves the final submission of your completed Written Project. This assessment makes up 50% of your overall module mark. Guidance: In your written project you should: Appropriately cover the topic of the extended project. Address the question(s) you set out to investigate, by researching, analysing and using information from relevant sources. Represent the complexities of the project topic. Analyse project outcomes and draw conclusions. Include a complete list of references used in producing the project. Please note: Kaplan International Pathways 1 kaplanpathways.com Both parts of this assessments are individual tasks, which means that you are expected to complete them by yourself. Before you submit your final written project, please make sure you carefully consider feedback you have received on your draft written project submission from your teacher. The suggestions made by your teacher will help you make your final submission more effective./nPart 2 Written Project You must choose a structure for your project which will best suit your topic and the question(s) you are investigating. Your project should be structured as follows: 1. Introduction (300 - 400 words) 2. Main part of project: exploration of sources through description of key issues/discussion/development/analysis Please note: this part of your project should have more than one section. The way you structure this should be decided by you, with support from your tutor. Your structure will depend on the topic you have chosen and the 'story you want to tell' (2350 words) 3. Conclusion (200-250 words) 4. References (not included in word count) 5. Appendices (if applicable) e.g.: any data sets you have used (not included in word count)/nPart 2 Written Project Your grade for Part 2 will contribute 50% of your overall module mark. It will be marked using the following areas and weightings: 5. Task achievement (25%) Have you followed the instructions for producing your Written Project? Have you chosen an appropriate topic for your project? Have you investigated the topic in sufficient detail? Have you addressed all the research questions? Have you considered the topic/questions from a variety of angles? Have you formatted your project appropriately and submitted it correctly, as specified in the instructions? 6. Use of sources (25%) Have you selected relevant sources for your project and used them in your project to support your claims? Have you synthesised information from a range of sources in your project? Have you analysed the information from the sources thoroughly and linked it with your project topic? Kaplan International Pathways 7. Structure (20%) Have you organised your project clearly into sections and arranged them in a logical order? Have you linked the ideas in your paragraphs? Do you use signposting language effectively? 8. Academic Integrity (15%) kaplanpathways.com Is your paraphrasing/summarising successful in terms of retention of meaning? Have you included an accurate citation with each paraphrase/summary? Are all quotes (if used) correctly marked? Do you have a full list of matching references which are correctly formatted (including organisation in alphabetical order and NO bullet points or numbering)? Are all facts/data presented genuine, accurate and precise? 9. Clarity of expression (15%) Have you checked the accuracy of your written language? Have you explained complicated concepts to help your reader? Have you used topic-specific vocabulary accurately? You will receive a % mark in each of these categories. The overall mark will be a percentage (0-100%)./n2 Introduction Over a century, important turning points in the history of aerospace innovation have transformed the skies above industry below. Presently, the industry is at the forefront of technological innovation, emphasizing the improvement of performance, efficiency, and safety through state-of-the-art innovations. The purpose of this report is to investigate how aerospace advancements can be applied to shape the future. We investigate how advances in materials science, manufacturing processes, and artificial intelligence (Al) are transforming the aerospace sector and beyond. Background Information on Aerospace Innovations There have been several significant historical tuming points in the intriguing journey of aerospace developments. The Wright brothers' successful powered flight in 1903 is considered the beginning of aviation. But it wasn't until the middle of the 20th century that the aerospace industry saw several notable breakthroughs that opened up new possibilities. Aerospace technology underwent a boom of improvements following the end of World War II. In order to increase speed and efficiency, military aircraft powered by piston engines were switched to jet engines during the conflict. This change paved the way for improvements in civilian aviation. In the 1950s, the military began developing small jet-powered aircraft for VIP transport. Learjet and Dassault Falcon were among the pioneering companies introducing private jets in the early 1960s, revolutionizing executive transportation. The introduction of the Boeing 747 in the 1970s marked a turning point in private and commercial air travel. With its expansive size and luxurious amenities, the 747 offered unprecedented comfort and convenience for private fliers, including heads of state and celebrities. Aerospace technology has undergone remarkable advancements over the decades, enhancing safety, efficiency, and comfort in private flying. The introduction of jet engines, which provide simplicity, dependability, and improved performance over piston engines, changed aircraft propulsion. The creation and broad application of GPS technology transformed aviation navigation. GPS replaced traditional ground- based navigation by improving aircraft safety, efficiency, and precision with pinpoint accuracy and real-time positional data Grady (2018)./nAerospace Innovations & Applications In order to improve performance, efficiency, and safety, the aerospace industry is at the forefront of technical innovation, continuously pushing the limits of materials and production processes. Two main areas of interest have surfaced in recent years: the investigation of lightweight materials for improved fuel efficiency and the use of additive manufacturing for quick prototyping and customization. Not only can lightweight materials save gasoline, but they also improve performance and payload capacity. High strength and stiffness CFRP composite materials are among the cutting-edge materials being investigated; these properties make them perfect for use in aeronautical applications. They are lighter than traditional materials like aluminum and offer excellent fatigue resistance. MMCs combine lightweight metals with ceramic reinforcements, resulting in materials that are both strong and lightweight. Intermetallic compounds offer superior strength and corrosion resistance compared to conventional alloys (Portolabs Manufacturing Accelerated 2021). The aircraft sector has experienced a change because of additive manufacturing (AM), sometimes referred to as 3D printing, which offers quick prototype and customization options. Engineers may design complicated interior structures and intricate geometries with additive manufacturing (AM) that are hard or impossible to accomplish with conventional manufacturing techniques. Engineers may test ideas and iterate designs more quickly with additive manufacturing (AM) before moving forward with full-scale production. Additionally, AM makes it possible to produce bespoke parts on demand that are suited to certain aircraft configurations or mission requirements. By leveraging lightweight materials and additive manufacturing technologies, aerospace companies can achieve significant improvements in fuel efficiency, performance, and operational flexibility, ensuring a sustainable and competitive future for the industry (Crawford, 2017). Autonomous drones represent a significant advancement in aviation technology, leveraging Al integration to achieve various objectives across industries. Autonomous drones equipped with Al integration are revolutionizing surveillance and security operations. These drones can autonomously patrol predefined areas, analyze live video feeds in real-time using computer vision algorithms, and identify potential security threats or anomalies. Al algorithms play a critical role in predictive/nmaintenance strategies with in the aviation industry, enabling proactive identification of potential equipment failures or maintenance issues before they occur. By analyzing vast amounts of sensor data and historical maintenance records, these algorithms can predict component failures, optimize maintenance schedules, and prevent costly unplanned downtime (Shah, Trivedi, & Tank, 2022). By detecting early warning signs such as abnormal vibrations, temperature fluctuations, or wear and tear, these algorithms can alert maintenance teams to take preemptive action, thereby reducing the likelihood of unscheduled maintenance events and optimizing asset reliability. Conclusion Aeronautics manufacturing faces significant challenges and ethical considerations. Ethical dilemmas emerge with Al integration, posing questions about accountability, and data privacy. The aerospace industry continues to soar on wings of innovation, driven by material advancements, manufacturing techniques & Al integration. From lightweight materials to additive manufacturing drones, each innovation reshapes the future of flight, promising enhanced efficiency, safety & sustainability.See Answer
Q15: 1. Introduction The Final Year Project (FYP) gives you a unique opportunity to study an area related to your course in depth. It also enables you to develop and demonstrate a range of transferable skills. The FYP is an important element of your Engineering degree programme. It represents 25% of the marks (30 credits) allocated for the 3rd year. Unlike other modules, the FYP cannot be compensated, so you must successfully pass in order to achieve an accredited Honours Degree. 1.1. Basis of the FYP Extended projects like the FYP permit the development of many of the generic skills that the graduate engineer may be expected to have. Included in this are: ● ● meeting deadlines and working within other externally defined constraints, tackling work which lacks a well-defined outcome, or which has a wide range of possible answers, ● opportunities to be creative, putting technical work in a social and commercial context, presentation of technical information in various ways, searching published sources of information, applying technical knowledge to unfamiliar problems. ● ● FINAL YEAR PROJECT ● ● 1.2. Objectives of the FYP In general, the FYP objectives are to demonstrate the ability to plan, execute and report on a significant, individual project under the guidance of a supervisor. In particular, the project objectives include: the planning and management of their own work over an extended period of time, ● ● ● To undertake practical, theoretical, simulation and/or design work in order to achieve a specified objective. To work on problems for which there may be no unique solution. To plan and manage a project in order to achieve the required objectives. To present your project process and outcomes in a professional manner. To individually and independently learn and gain knowledge and experience in a specified subject area. 1.3. About this document This Handbook document represents your FYP guide for the 2022-2023 Academic Year. It contains information concerning the running and assessment of the project { ² } 2 FINAL YEAR PROJECT which will help you to successfully complete your project. Please read this document carefully, and consult its contents as often as required throughout the duration of the project. ( ³ ) 2. General Structure The FYP is carried out in two parts, as part of two different modules: Part 1 (Trimester 1), worth 10 credits, which represents 50% the Industrial Management and Project Preparation module. Part 2 (Trimester 2), worth 20 credits, which is a standalone module. ● ● FINAL YEAR PROJECT Marks you receive for Part 1 represent 50% of the final mark for the Industrial Management and Project Preparation module. Marks you receive for Part 2 represent 100% of the Level 3.2 Project module. There is no final, aggregate mark for the FYP but you will work on the same project topic throughout Trimesters 1 and 2. Trimester 1 Industrial Management and Project Preparation module Industrial Management FYP Part 1 (50%) component (50%) Module mark: 50% Industrial Management 50% FYP Part 1 (10 credits + 10 credits) {4} Trimester 2 Level 3.2 Project module FYP Part 2 (100%) Module mark: 100% FYP Part 2 (20 credits) FINAL YEAR PROJECT 3. Assessment and Deadlines 3.1. Part 1 submission and deadline For Part 1 of the project, you will need to prepare and submit an electronic copy of a report. The structure and content of the Part 1 report are detailed in Section 8 of the Handbook. The submission of the electronic copy of the report (in PDF or Word formats only) will be made on Blackboard, via a portal on the Industrial Management and Project Preparation module page. 3.2. Part 2 submission and deadline For Part 2 of the project, you will need to prepare and submit an electronic copy of a report, as well as prepare and deliver a 15-minute presentation. The structure and content of the Part 2 report are detailed in Section 9 of the Handbook, while guidance on how to prepare the presentation is provided in Section 12 of the Handbook. The submission of the electronic copy of the report (in PDF or Word formats only) will be made on Blackboard, via a portal on the Level 3.2 Project module page. 3.3. Breakdown of assessment marks Your Part 1 and 2 reports, as well as the Part 2 presentation will be assessed not only by your supervisor, but also by an independent moderator. The moderator is an academic who is not familiar with your project and who will be able to provide a second, objective evaluation of your academic performance. The following table summarises the breakdown of assessment marks. Note that the marks awarded by your supervisor and the independent moderator carry equal weight. The Part 1 mark is 100% determined based on the report. The Part 2 mark is determined 90% based on the report mark and 10% based on the presentation mark. The detailed marking schemes your supervisor and moderator will use to assess your work are detailed in Appendices F, G and H of the Handbook. {5} ● ● ● ● Part ● Part 1 Part 2 Credits 10 credits 20 credits FINAL YEAR PROJECT Assessment Type and Weighting Report 100% Report 90% Presentation 10% Marks Given By Supervisor 50% Moderator 50% Supervisor 50% Moderator 50% 3.4. Late submissions, non-submissions, and RAPS The FYP submissions are subject to the university's standard penalties for late submission: Joint mark by supervisor and moderator Overall 50% of Industrial Management module mark 100% of Level 3.2 Project module mark The late submission period is 7 consecutive days following the original deadline. If the work is submitted within this 7-day late submission period, and the work is at least of pass standard, then the mark is capped at 40%. If the work is below pass standard, then no late submission penalties apply. If you have an accepted PMC for late submission, then the late submission penalty is not applied. You must ensure you submit before the 7 days have passed, or your work will be recorded as a non-submission (see next point). If the work is submitted more than 7 days late, then it will be recorded as a non- submission and no marks will be awarded. If you know you will submit more than 7 days late (due to various reasons), you will need to get an accepted PMC for non-submission. Applying for a non- submission PMC must be carefully considered however, because you will only be allowed to submit again during the resit period (see point 3.6 below). This means that you will graduate late. For Part 2 of the project, in which you submit both a report and deliver a presentation, the above points apply to each component independently (e.g., you can submit your report on time, but fail to attend the presentation, and the presentation will be recorded as a non-submission). If you have a Reasonable Adjustment Plan (RAP) or Carer Plan (CP) which grants you an extra calendar week, this applies the FYP as well. All your submission deadlines will be one week later compared to what is officially published. Since with a RAP or CP your deadline is extended by one calendar week, the late submission period is 7 consecutive days following the extended deadline. VERY IMPORTANT NOTE: as part of your project, you will be expected to think on how to deal with unexpected problems (e.g., equipment failure, software problems, etc.) and have some mitigation strategies in place (see Section 6.2 of {6}See Answer
Q16:2. Using the website, airfoiltools.com, draw NACA 4420. Show your drawing process.See Answer
Q17:Question 11. An aircraft with a stall speed of V m s¹ in straight, steady and level flight performs a horizontal turn of radius R m. What is the slowest speed at which the turn can be performed if the turn is limited by stall? The answer should have units of m s.¹. [4 Marks]See Answer
Q18:Question 8. A jet engine aircraft is required to cruise for 3000 km at a constant altitude of 8 km. At the end of cruise the aircraft will weigh W KN. The wing area of the aircraft is 80 m², Co for the flight is 0.04, CL is 0.3 and the aircraft's thrust specific fuel consumption is 105 kg N-¹ s¹. How much fuel will be used in the cruise? The answer should have units of KN. [4 Marks]See Answer

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