from __future__ import division # let 5/2 = 2.5 rather than 2
from Aerothon.ACPropeller import ACPropeller
from Aerothon.AeroUtil import STDCorrection
import numpy as npy
import pylab as pyl
from scalar.units import IN, LBF, SEC, ARCDEG, FT, RPM, OZF, GRAM, gacc, hPa, K, W, inHg
from scalar.units import AsUnit
# Set Propeller properties
Prop = ACPropeller()
Prop.name = 'APC 12.2x4.5'
Prop.D = 12.2 * IN
Prop.Thickness = 5 / 8 * IN
#Prop.PitchAngle = 12*ARCDEG
Prop.Pitch = 4.5 * IN
Prop.dAlpha = 3.4 * ARCDEG
Prop.Solidity = 0.0135
Prop.AlphaStall = 14 * ARCDEG
Prop.CLSlope = 0.072 / ARCDEG
Prop.Weight = 1.80 * OZF
#
# These are corrected for standard day
#Standard correction for 2:00 pm for the test day
STD = STDCorrection(30.03 * inHg, (19 + 273.15) * K)
#
# RPM, Thrust
Prop.ThrustData = [(3750 * RPM, (0 * LBF + 14 * OZF) * STD),
(6210 * RPM, (2 * LBF + 4 * OZF) * STD),
(7830 * RPM, (1 * LBF + 15 * OZF) * STD),
# import Aerothon modules
from Aerothon.scalar.units import IN, LBF, SEC, ARCDEG, FT, RPM, OZF, GRAM, gacc, W, K,\
degR, inHg, MM
from Aerothon.scalar.units import AsUnit
from Aerothon.ACPropeller import ACPropeller
from Aerothon.AeroUtil import STDCorrection
#==============================================================================#
# PROPELLER MODEL
#==============================================================================#
# Set Propeller properties
Prop = ACPropeller()
Prop.name = 'APC 20x8E'
Prop.D = 20*IN
Prop.Thickness = 0.5*IN
Prop.Pitch = 8*IN
Prop.dAlpha = 3.3*ARCDEG
Prop.Solidity = 0.0126
Prop.AlphaStall = 20*ARCDEG
Prop.AlphaZeroCL = 0*ARCDEG
Prop.CLSlope = ..2/ARCDEG #- 2D airfoil lift slope
Prop.CDCurve = 2.2 #- 2D curvature of the airfoil drag bucket
Prop.CDp = .02 #- Parasitic drag
Prop.Weight = 4.05*OZF
Prop.ThrustUnit = LBF
Prop.ThrustUnitName = 'lbf'
from __future__ import division # let 5/2 = 2.5 rather than 2 #from os import environ as _environ; _environ["scalar_off"] = "off" from Aerothon.ACPropeller import ACPropeller from Aerothon.ACEngine import ACEngine from Aerothon.ACPropulsion import ACPropulsion import numpy as npy from scalar.units import IN, LBF, PSFC, SEC, ARCDEG, FT, OZF, RPM, HP from scalar.units import AsUnit # Set Propeller properties Prop = ACPropeller() Prop.D = 14.5 * IN Prop.Thickness = .5 * IN Prop.PitchAngle = 12 * ARCDEG Prop.dAlpha = 0 * ARCDEG Prop.Solidity = 0.0136 Prop.RD = 3 / 8 Prop.AlphaStall = 14 * ARCDEG Prop.Weight = 3 / 32 * LBF # Set Engine properties Engine = ACEngine() Engine.Rbs = 1.1 Engine.Rla = 3.5 Engine.NumCyl = 1 Engine.NumRev = 1 Engine.CompRatio = 9 Engine.Vd = 0.607 * IN**3 Engine.PistonSpeedR = 38.27 * FT / SEC
from __future__ import division # let 5/2 = 2.5 rather than 2 from Aerothon.ACPropeller import ACPropeller from Aerothon.ACEngine import ACEngine from Aerothon.ACPropulsion import ACPropulsion import numpy as npy from scalar.units import IN, LBF, PSFC, SEC, ARCDEG, FT, OZF, RPM, HP from scalar.units import AsUnit # Set Propeller properties Prop = ACPropeller() Prop.D = 13.5 * IN # Diameter Prop.Thickness = .5 * IN # Thickness at the hub... just for drawing purposes Prop.PitchAngle = 14 * ARCDEG # Pitch angle.. aka \Beta Prop.dAlpha = 0 * ARCDEG # Difference between measured alpha and zero lift alpha Prop.Solidity = 0.0136 # Proportional to the blade disk area, similar to the activity factor (AreaBlades/(2*D**2)) Prop.RD = 3 / 8 # The location on the profile chord where the PitchAngle is defined (default 3/8) Von Mises 306 Prop.AlphaStall = 12 * ARCDEG # Stall angle of attack Prop.Weight = 3 / 32 * LBF # Weight # Use these parameters to match test data if need be. # Prop.CLSlope = .078/ARCDEG #- 2D airfoil lift slope # Prop.CDCurve = 2.2 #- 2D curvature of the airfoil drag bucket # Prop.CDp = .02 #- Parasitic drag # Set Engine properties - glow engine... see 2015 files for setting up electric motor Engine = ACEngine() Engine.Rbs = 1.1 Engine.Rla = 3.5 Engine.NumCyl = 1 Engine.NumRev = 1 Engine.CompRatio = 9
