1 / 39

Research & Development

Research & Development. Wind Tunnel Testing Needs to be applicable for subsonic, transonic, and supersonic velocities Scaled down model Associated costs Dimensions to minimize error due to scale effects, flow blockage, and wall boundary layers Possible testing locations

dwilson
Download Presentation

Research & Development

An Image/Link below is provided (as is) to download presentation Download Policy: Content on the Website is provided to you AS IS for your information and personal use and may not be sold / licensed / shared on other websites without getting consent from its author. Content is provided to you AS IS for your information and personal use only. Download presentation by click this link. While downloading, if for some reason you are not able to download a presentation, the publisher may have deleted the file from their server. During download, if you can't get a presentation, the file might be deleted by the publisher.

E N D

Presentation Transcript

  1. Research & Development • Wind Tunnel Testing • Needs to be applicable for subsonic, transonic, and supersonic velocities • Scaled down model • Associated costs • Dimensions to minimize error due to scale effects, flow blockage, and wall boundary layers • Possible testing locations • NASA Glenn Research Center – Cleveland, OH • NASA Langley Research Center - Hampton, VA • Purdue’s Mach-6 supersonic tunnel – West Lafayette, IN AAE 450 Spring 2008 Aerothermal 1

  2. Assumptions: • Used Historical Values for large variety of similar shaped rockets and scaled the drag coefficient accordingly to determine CD at α=0. • Also attempted CFD to determine CD at α=0. • Then used CD at α=0 in order to generate plots of CD versus AoA. • -”Normal” Geometry indicates all upper stages are smaller in diameter than their predeceasing lower stage and only a total of 1 to 2 shoulders. AAE 450 Spring 2008 Aerothermal 2

  3. -Used pressure coefficient to calculate the axial and normal force coefficients. -Used the axial and normal force coefficients to calculate the drag coefficient. AAE 450 Spring 2008 Aerothermal 3

  4. Coefficient of Drag Plot Authors: Woods, Zott AAE 450 Spring 2008 Aerothermal 2

  5. Vanguard Check – subsonic case using Fluent Velocity VectorsRed – 401 m/sGreen – 225 m/s Static PressureRed – 1.18 atmBlue – .364 atm AAE 450 Spring 2008 Aerothermal 5

  6. Vanguard Check – subsonic case using Fluent Static TemperatureRed – 300 KBlue – 220 K AAE 450 Spring 2008 Aerothermal 6

  7. 1 kg launch vehicle – Mach 1 case using Fluent PressureRed – 1.56 atmBlue – .373 atm AAE 450 Spring 2008 Aerothermal 7

  8. 1 kg launch vehicle – Mach 1 case using Fluent Velocity Red – 411 m/s Green – 208 m/s AAE 450 Spring 2008 Aerothermal 8

  9. 200 g Aerodynamic Loads AAE 450 Spring 2008 Aerothermal 9

  10. 1 kg Aerodynamic Loads AAE 450 Spring 2008 Aerothermal 10

  11. 5 kg Aerodynamic Loads AAE 450 Spring 2008 Aerothermal 11

  12. Creation of the Pressure Distribution AAE 450 Spring 2008 Aerothermal 12

  13. Linear Perturbation Forces are found by an integration of pressure distribution over the launch vehicle exterior Integrations are done numerically within the code Phi is the geometric angle w/respect to the freestream S is a reference area, taken to be the area of the base of the launch vehicle Validated by comparison to Vanguard results and other related geometries AAE 450 Spring 2008 Aerothermal 13

  14. Important Assumptions in Theory Small changes in geometry Small angle of attack ~ 0 – 14 degrees Valid for subsonic or supersonic flow 0 < M < 0.88 1.12 < M < 5 Axial force neglects viscous effects AAE 450 Spring 2008 Aerothermal 14

  15. Stresses due to Aerodynamic Force Shear Stress Differential of Normal Force between stages Axial Loading A*q∞*S Bending Moment Picture by Jayme Zott and Alex Woods AAE 450 Spring 2008 Aerothermal 15

  16. Pictures AAE 450 Spring 2008 Aerothermal 16

  17. Pictures AAE 450 Spring 2008 Aerothermal 17

  18. Heating Rate • Heating Rate Analysis • Primarily done to design a TPS (Thermal Protection System) • Stagnation Point • Theoretical analysis using methods outlined by Professor Schneider • Nose cone heating • Determine best material and thickness for the structure of the nose cone • Alternative methods and materials • SODDIT (Sandia One-Dimensional Direct and Inverse Thermal) • Ablative materials AAE 450 Spring 2008 Aerothermal 18

  19. Heating Rate 3 0 1 2 • Assumptions • Constant specific heats • No heat transfer within the body • Treat whole nosetip as one solid heat sink • Laminar flow at point 2 • No convective heating at point 3, only radiative • Wall temperature is the same at all 4 points Lumped Heating at Solid Nosetip Method AAE 450 Spring 2008 Aerothermal 19

  20. Equations M = 3, N = 0.5 for fully catalytic surface M2 = 3.2 for laminar, flat plate = nose body radius = radiation from fluid to surface = volume of solid nosetip Heating rate at point 2 on the nose cone AAE 450 Spring 2008 Aerothermal 20

  21. Heating Rate Matlab code: *help from Vince Teixeira AAE450_Stag_heat_analysis.m Uses trajectory outputs (.mat files) Input: d - diameter (m) v - velocity (m/s) r - position from the center of the earth (m) c_p - specific heat of material (J/kg*K) rho_w - density of material (kg/m3) emiss - emissivity of material Output: q_dot - heating rate (W/m2) tw - thickness (mm) Tw - wall temperature (K) AAE 450 Spring 2008 Aerothermal 21

  22. Heating Rate AAE 450 Spring 2008 Aerothermal 22

  23. Backup Slides- Sizing Code Tables • Initial Sizing Code Table of Results Table Created by Chris Strauss AAE 450 Spring 2008 Aerothermal 23

  24. Backup Slides-CFD • Models to be used for GAMBIT griding of project rocket • Initial models of project rocket • Model would have been used to simulate each stage of flight in Fluent Models Created by Chris Strauss AAE 450 Spring 2008 Aerothermal 24

  25. Backup Slides-CFD • CMARC Model • Model of aircraft launched rocket initially conceived • Model was flexible enough so that multiple configurations could be made quickly • Model was scrapped after it was discovered CMARC results are only valid to Mach 0.9 Model Created by Chris Strauss AAE 450 Spring 2008 Aerothermal 25

  26. Drag Coefficient Standard Deviation • Method • Create a randomizer that produces random values of angle of attack from 0-10 degrees • Fed angles of attack into Cd code to obtain values for Cd • Cd code created by Jayme Zott • Entered values for Cd into Excel to calculate standard deviation with standard deviation function AAE 450 Spring 2008 Aerothermal 26

  27. Wing Moment Coefficient versus AoA Fig. by Brian Budzinski Top Down View Fig. by Kyle Donohue Though an aircraft launch was not put into operation. A wing would be beneficial if it were. A wing creates an additional nose up pitching moment allowing the launch vehicle to pitch from an initial horizontal configuration (α=0°) into a final vertical configuration (α=90°). AAE 450 Spring 2008 Aerothermal 27

  28. Shear on Launch Vehicle from Wing Fig. by Brian Budzinski Shear Coefficient on Launch Vehicle from Wing Fig. by Brian Budzinski The shear created through the addition of a wing or fins is assumed to be equal to the normal force caused by the corresponding part. Shear on Launch Vehicle from Fins Fig. by Brian Budzinski AAE 450 Spring 2008 Aerothermal 28

  29. Wing Axial Force Coefficient versus AoA Fig. by Brian Budzinski Wing Normal Force Coefficient versus AoA Fig. by Brian Budzinski ASSUMPTIONS: Initial Horizontal Launch Configuration Final Vertical Configuration Newtonian Model Delta Wing AAE 450 Spring 2008 Aerothermal 29

  30. Wing Lift Coefficient versus AoA Fig. by Brian Budzinski Wing Drag Coefficient versus AoA Fig. by Brian Budzinski ASSUMPTIONS: Initial Horizontal Launch Configuration Final Vertical Configuration Newtonian Model Delta Wing AAE 450 Spring 2008 Aerothermal 30

  31. Once the lift and drag coefficients are determined, the lift versus drag curve can be created. Drag Coefficient versus Lift Coefficient Fig. by Brian Budzinski ASSUMPTIONS: Initial Horizontal Launch Configuration Final Vertical Configuration Newtonian Model Delta Wing AAE 450 Spring 2008 Aerothermal 31

  32. Launch vehicle with a pair of fins. • Beneficial for: • Stability Control • Ground Launch • Aircraft Launch • Balloon Launch Side View Fig. by Kyle Donohue Fins were not implemented because D&C was able to successfully control the launch vehicle without them. AAE 450 Spring 2008 Aerothermal 32

  33. Wing Analysis Divide the wing up into two sections: leading edge and lower surface. These two are chosen because they are the two portions exposed to the relative wind once given an angle of attack. AAE 450 Spring 2008 Aerothermal 33

  34. Wing AnalysisContinued Lower Surface Eqns. A similar analysis can be done for a pair of fins. AAE 450 Spring 2008 Aerothermal 34

  35. References Ashley, Holt, Engineering Analysis of Flight Vehicles, Dover Publications Inc., New York, 1974, pp. 303-312 Anderson, John D., Fundamentals of Aerodynamics, Mcgraw-Hill Higher Education, 2001 Professor Colicott, in reference to linearized theory applications AAE 450 Spring 2008 Aerothermal 35

  36. References Barrowman, James and Barrowman, Judith, "The Theoretical Prediction of the Center of Pressure" A NARAM 8, August 18, 1966. www.Apogeerockets.com Klawans, B. and Baughards, J. "The Vanguard Satellite Launching Vehicle - an engineering summary" Report No. 11022, April 1960 Morrisette, E. L., Romeo D. J., “Aerodynamic Characteristics of a Family of Multistage Vehicles at a Mach Number of 6.0”, NASA TN D-2853, June 1965 Professor Williams, concerning the use of pressure coefficients to determine aerodynamic forces The entire Aerothermodynamics group for their invaluable help and support

  37. References • Anderson Jr., John D., “Hypersonic and High-Temperature Gas Dynamics”, 2nd ed., AIAA, Reston, VA, 2006. • Schneider, Steven P., “Methods for Analysis of Preliminary Spacecraft Designs”, AAE450, Spacecraft Design, Purdue University • Schneider, Steven P., personal conversation • http://www.omega.com/literature/transactions/volume1/emissivitya.html • http://ntrs.nasa.gov/archive/nasa/casi.ntrs.nasa.gov/20080005052_2008005139.pdf • http://ntrs.nasa.gov/archive/nasa/casi.ntrs.nasa.gov/19740009531_1974009531.pdf

  38. References • Wade, M., “Vanguard”, 1997-2007. [http://www.astronautix.com/lvs/vanguard.htm] • Tsohas, J., “AAE 450 Spacecraft Design Spring 2008: Guest Lecture Space Launch Vehicle Design”, 2008 • “The Vanguard Report”, The Martin Company, Engineering Report No. 11022, April 1960 AAE 450 Spring 2008 Aerothermal 38

  39. References • Hankey, Wilbur L., Re-Entry Aerodynamics, AIAA, Washington D.C., 1988, pp. 70-73 • Rhode, M.N., Engelund, W.C., and Mendenhall, M.R., “Experimental Aerodynamic Characteristics of the Pegasus Air-Launched Booster and Comparisons with Predicted and Flight Results”, AIAA Paper 95-1830, June 1995. • Anderson, John D., Fundamentals of Aerodynamics, Mcgraw-Hill Higher Education, 2001 • Ashley, Holt, Engineering Analysis of Flight Vehicles, Dover Publications Inc., New York, 1974, pp. 303-312 • The Martin Company, “The Vanguard Satellite Launching Vehicle”, Engineering Report No. 11022, April 1960. AAE 450 Spring 2008 Aerothermal

More Related