def calc(self):
'''
Calculates the landing field length from the wing loading and the maximum lift at landing.
All calculations are on sea level
:Source: Airplane Design Part I, J. Roskam, DARCorporation, 2005, Fourth Edition, p.111, eq. 3.15-3.16
'''
wsL = self.parent.wsL.getValue()
cLL = self.parent.cLL.getValue()
rhoAP = self.parent.atmosphere.rhoAP.getValue()
#Stall speed Calculation
Vs = sqrt(2 * wsL * 9.81 / (rhoAP * cLL)) #Eq. 3.1 from Roskam I p-90
#Approach Speed Calculation
Va = 1.3 * Vs * 3.6
self.parent.vAPPR = parameter(Va, name='vAPPR', parent=self.parent)
#make the Roskam Assumption
#Roskam assumes ft and kts
#kts to m is 1,8**2
return self.setValueCalc(0.3048 * 0.3 * 0.539956803 ** 2 * Va ** 2)
###################################################################################################
#EOFEOFEOFEOFEOFEOFEOFEOFEOFEOFEOFEOFEOFEOFEOFEOFEOFEOFEOFEOFEOFEOFEOFEOFEOFEOFEOFEOFEOFEOFEOFEOFE#
###################################################################################################
def cpacsExport(self, CPACSObj):
'''
this methods exports all parameters nested in the component. Nested Components will be called as well.
cpacsPath must be filled
'''
def exportVector(CPACSObj, cpacsPath, value):
myVector = getObjfromXpath(CPACSObj, cpacsPath)
myType = "vector"
myVector.set_valueOf_(str(value))
myVector.set_mapType(myType)
exportVector(
CPACSObj,
'/cpacs/vehicles/aircraft/model/analyses/aeroPerformanceMap/machNumber',
self.aircraft.atmosphere.MaCR.getValue())
exportVector(
CPACSObj,
'/cpacs/vehicles/aircraft/model/analyses/aeroPerformanceMap/angleOfAttack',
'0;2;4')
exportVector(
CPACSObj,
'/cpacs/vehicles/aircraft/model/analyses/aeroPerformanceMap/angleOfYaw',
'0')
exportVector(
CPACSObj,
'/cpacs/vehicles/aircraft/model/analyses/aeroPerformanceMap/reynoldsNumber',
self.aircraft.wing.reynoldsNr.getValue())
self.lilipolint = parameter(
cpacsPath=
'/cpacs/toolspecific/liftingLine/toolParameters/usePOLINT',
value="false")
super(tool, self).cpacsExport(CPACSObj)
def cpacsExport(self, CPACSObj):
'''
this methods exports all parameters nested in the component. Nested Components will be called as well.
cpacsPath must be filled
'''
def exportVector(CPACSObj, cpacsPath, value):
myVector = getObjfromXpath(CPACSObj, cpacsPath)
myType = "vector"
myVector.set_valueOf_(str(value))
myVector.set_mapType(myType)
exportVector(CPACSObj, '/cpacs/vehicles/aircraft/model/analyses/aeroPerformanceMap/machNumber',
self.aircraft.atmosphere.MaCR.getValue())
exportVector(CPACSObj, '/cpacs/vehicles/aircraft/model/analyses/aeroPerformanceMap/angleOfAttack', '0;2;4')
exportVector(CPACSObj, '/cpacs/vehicles/aircraft/model/analyses/aeroPerformanceMap/angleOfYaw', '0')
exportVector(CPACSObj, '/cpacs/vehicles/aircraft/model/analyses/aeroPerformanceMap/reynoldsNumber',
self.aircraft.wing.reynoldsNr.getValue())
self.lilipolint = parameter(cpacsPath='/cpacs/toolspecific/liftingLine/toolParameters/usePOLINT', value="false")
super(tool, self).cpacsExport(CPACSObj)
def calc(self):
'''
Calculates the landing field length from the wing loading and the maximum lift at landing.
All calculations are on sea level
:Source: Airplane Design Part I, J. Roskam, DARCorporation, 2005, Fourth Edition, p.111, eq. 3.15-3.16
'''
wsL = self.parent.wsL.getValue()
cLL = self.parent.cLL.getValue()
rhoAP = self.parent.atmosphere.rhoAP.getValue()
#Stall speed Calculation
Vs = sqrt(2 * wsL * 9.81 / (rhoAP * cLL)) #Eq. 3.1 from Roskam I p-90
#Approach Speed Calculation
Va = 1.3 * Vs * 3.6
self.parent.vAPPR = parameter(Va, name='vAPPR', parent=self.parent)
#make the Roskam Assumption
#Roskam assumes ft and kts
#kts to m is 1,8**2
return self.setValueCalc(0.3048 * 0.3 * 0.539956803**2 * Va**2)
def __init__(self, aircraft):
'''
will initialize instance mainly used for documentation
links to the aircraft instance
initiates the airfoil class
@Method: Component Constructor
'''
component.__init__(self)
self.id = 'vtp'
self.aircraft = aircraft
self.level = 2
self.UID = uID(cpacsPath='/cpacs/toolspecific/vampZero/vtpUID')
# Mass
self.mVtp = mVtp(
cpacsPath=
'/cpacs/vehicles/aircraft/model/analyses/massBreakdown/mOEM/mEM/mStructure/mWingsStructure/mWingStructure[3]/massDescription['
+ self.id + '_mass]/mass')
self.parentUID = parameter(
value=self.id,
cpacsPath=
'/cpacs/vehicles/aircraft/model/analyses/massBreakdown/mOEM/mEM/mStructure/mWingsStructure/mWingStructure[3]/massDescription['
+ self.id + '_mass]/parentUID')
# CoG
self.posCoG = posCoG(
cpacsPath=
'/cpacs/vehicles/aircraft/model/analyses/massBreakdown/mOEM/mEM/mStructure/mWingsStructure/mWingStructure[3]/massDescription['
+ self.id + '_mass]/location/x')
# Geometry
self.refArea = refArea()
self.refAreaTakeOff = refAreaTakeOff()
self.refAreaLanding = refAreaLanding()
self.expArea = expArea()
self.wetArea = wetArea()
self.aspectRatio = aspectRatio()
self.span = span()
self.taperRatio = taperRatio()
self.tcAVG = tcAVG()
self.cRoot = cRoot()
self.cTip = cTip()
self.cMAC = cMAC()
self.yMAC = yMAC()
self.xMAC = xMAC()
self.phiLE = phiLE()
self.phiTE = phiTE()
self.phi25 = phi25()
self.phi50 = phi50()
self.xRoot = xRoot()
self.zRoot = zRoot()
self.lVT = lVT()
self.cVT = cVT()
self.LoD = LoD()
# Aerodynamics
self.cLalpha = cLalpha()
self.cLMAX = cLMAX()
self.formFactor = formFactor()
self.reynoldsNr = reynoldsNr()
self.cfLAM = cfLAM()
self.cfTURB = cfTURB()
self.cD0c = cD0c()
self.nLam = nLam()
# Airfoil
self.airfoilr = airfoil(self) # Root Airfoil
self.airfoilt = airfoil(self) # Tip Airfoil
# Rudder
self.rudder = rudder(self)
def __init__(self):
'''
will initialize all values upon creation of the instance. Mainly used for documentation
'''
component.__init__(self)
self.id = 'aircraft'
self.level = 1
self.modelUID = uID(cpacsPath='/cpacs/toolspecific/vampZero/aircraftModelUID')
self.version = parameter(value=0.1, cpacsPath='/cpacs/header/version')
# Mass Data
self.mLM = mLM(cpacsPath='/cpacs/vehicles/aircraft/model/analyses/massBreakdown/designMasses/mMLM/mass')
self.mTOM = mTOM(cpacsPath='/cpacs/vehicles/aircraft/model/analyses/massBreakdown/designMasses/mTOM/mass')
self.mZFW = mZFW(cpacsPath='/cpacs/vehicles/aircraft/model/analyses/massBreakdown/designMasses/mZFM/mass')
self.oEM = oEM(cpacsPath='/cpacs/vehicles/aircraft/model/analyses/massBreakdown/mOEM/massDescription/mass')
self.oIM = oIM(
cpacsPath='/cpacs/vehicles/aircraft/model/analyses/massBreakdown/mOEM/mOperatorItems/massDescription/mass')
# CoG
self.posCoG = posCoG(
cpacsPath='/cpacs/vehicles/aircraft/model/analyses/massBreakdown/designMasses/mTOM/location/x')
self.posNP = posNP()
self.posCoGOEM = posCoGOEM()
self.posCoGMIN = posCoGMIN()
self.posCoGMAX = posCoGMAX()
self.static_margin = static_margin()
# Inertia
self.massIXmTOM = massIXmTOM(
cpacsPath='/cpacs/vehicles/aircraft/model/analyses/massBreakdown/designMasses/mTOM/massInertia/Jxx')
self.massIYmTOM = massIYmTOM(
cpacsPath='/cpacs/vehicles/aircraft/model/analyses/massBreakdown/designMasses/mTOM/massInertia/Jyy')
self.massIZmTOM = massIZmTOM(
cpacsPath='/cpacs/vehicles/aircraft/model/analyses/massBreakdown/designMasses/mTOM/massInertia/Jzz')
self.massIXmZFW = massIXmZFW(
cpacsPath='/cpacs/vehicles/aircraft/model/analyses/massBreakdown/designMasses/mZFM/massInertia/Jxx')
self.massIYmZFW = massIYmZFW(
cpacsPath='/cpacs/vehicles/aircraft/model/analyses/massBreakdown/designMasses/mZFM/massInertia/Jyy')
self.massIZmZFW = massIZmZFW(
cpacsPath='/cpacs/vehicles/aircraft/model/analyses/massBreakdown/designMasses/mZFM/massInertia/Jzz')
self.massIXoEM = massIXoEM(
cpacsPath='/cpacs/vehicles/aircraft/model/analyses/massBreakdown/mOEM/massDescription/massInertia/Jxx')
self.massIYoEM = massIYoEM(
cpacsPath='/cpacs/vehicles/aircraft/model/analyses/massBreakdown/mOEM/massDescription/massInertia/Jyy')
self.massIZoEM = massIZoEM(
cpacsPath='/cpacs/vehicles/aircraft/model/analyses/massBreakdown/mOEM/massDescription/massInertia/Jzz')
# Performance Data
self.desRange = desRange() #@todo: set correct xpath (cpacsPath='/cpacs/vehicles/aircraft/model/global/designRange')
self.machCR = machCR(cpacsPath='/cpacs/vehicles/aircraft/model/global/machCruise')
self.altCR = altCR()
self.sTOFL = sTOFL()
self.sTOFLISA = sTOFLISA()
self.sLFL = sLFL() #@todo: set correct xpath (cpacsPath='/cpacs/vehicles/aircraft/model/global/sFL')
self.rangeType = rangeType()
self.timeRES = timeRES()
self.timeDESCENT = timeDESCENT()
self.timeCLIMB = timeCLIMB()
self.timeCR = timeCR()
self.distRES = distRES()
self.distDESCENT = distDESCENT()
self.distCLIMB = distCLIMB()
self.distCR = distCR()
self.gammaCLIMB = gammaCLIMB()
self.gammaDESCENT = gammaDESCENT()
self.IASCLIMB = IASCLIMB()
self.IASDESCENT = IASDESCENT()
# Crew Data
self.nPilot = nPilot()
self.nCabinCrew = nCabinCrew()
# Sizing Data
self.wsTO = wsTO()
self.wsL = wsL()
self.wsMAX = wsMAX()
self.twTO = twTO()
self.twFAR20121a = twFAR25121a()
self.twFAR20121b = twFAR25121b()
self.twTOP25 = twTOP25()
# Aerodynamics Data
self.cLCR = cLCR()
self.cLMAX = cLMAX()
self.cLTO = cLTO()
self.cLL = cLL()
self.cDCR = cDCR()
self.cDTO = cDTO()
self.cDL = cDL()
self.cD0 = cD0()
self.loDCR = loDCR()
self.loDTO = loDTO()
self.loDLOI = loDLOI()
self.loDCLIMB = loDCLIMB()
self.oswald = oswald()
self.cLROLL = cLROLL()
# DOC
self.aFactor = aFactor()
self.inflation = inflation()
self.USDexchangeEURO = USDexchangeEURO()
self.tBlock = tBlock()
self.tFlight = tFlight()
self.utilization = utilization()
self.costDepreciation = costDepreciation()
self.costFuel = costFuel()
self.costMaintenance = costMaintenance()
self.costAircraftMaintenance = costAircraftMaintenance()
self.costEngineMaintenance = costEngineMaintenance()
self.costCrew = costCrew()
self.costCap = costCap()
self.costGround = costGround()
self.costNavigation = costNavigation()
self.costLanding = costLanding()
self.costInterest = costInterest()
self.costInsurance = costInsurance()
self.costEmissionTrade = costEmissionTrade()
self.flightCycles = flightCycles()
self.priceFuel = priceFuel()
self.priceAircraft = priceAircraft()
self.DOC = DOC()
self.COC = COC()
self.COO = COO()
self.C1 = C1()
self.C2 = C2()
self.aFactor = aFactor()
self.costCap = costCap()
self.flightCycles = flightCycles()
#Stability
#self.staticMargin = staticMargin()
#Controllability
#self.refAreaHtpMIN = refAreaHtpMIN()
# Component Objects
self.wing = wing(self)
self.strut = strut(self)
self.fuselage = fuselage(self)
self.vtp = vtp(self)
self.htp = htp(self)
self.engine = engine(self)
self.landingGear = landingGear(self)
self.systems = systems(self)
self.payload = payload(self)
self.pylon = pylon(self)
self.fuel = fuel(self)
self.atmosphere = atmosphere(self)
self.tool = tool(self)
def __init__(self, aircraft):
'''
will initialize instance mainly used for documentation
links to the aircraft instance
initiates the airfoil class
'''
component.__init__(self)
self.id = 'htp'
self.aircraft = aircraft
self.level = 2
self.UID = uID(cpacsPath='/cpacs/toolspecific/vampZero/htpUID')
# Mass
self.mHtp = mHtp(
cpacsPath='/cpacs/vehicles/aircraft/model/analyses/massBreakdown/mOEM/mEM/mStructure/mWingsStructure/mWingStructure[2]/massDescription['+self.id+'_mass]/mass')
self.parentUID = parameter(value=self.id,
cpacsPath='/cpacs/vehicles/aircraft/model/analyses/massBreakdown/mOEM/mEM/mStructure/mWingsStructure/mWingStructure[2]/massDescription['+self.id+'_mass]/parentUID')
# CoG
self.posCoG = posCoG(
cpacsPath='/cpacs/vehicles/aircraft/model/analyses/massBreakdown/mOEM/mEM/mStructure/mWingsStructure/mWingStructure[2]/massDescription['+self.id+'_mass]/location/x')
# Geometry
self.refArea = refArea()
self.refAreaTakeOff = refAreaTakeOff()
self.refAreaCruise = refAreaCruise()
self.refAreaLanding = refAreaLanding()
self.expArea = expArea()
self.wetArea = wetArea()
self.aspectRatio = aspectRatio()
self.span = span()
self.taperRatio = taperRatio()
self.tcAVG = tcAVG()
self.cRoot = cRoot()
self.cTip = cTip()
self.cMAC = cMAC()
self.yMAC = yMAC()
self.xMAC = xMAC()
self.xMAC25 = xMAC25()
self.phiLE = phiLE()
self.phiTE = phiTE()
self.phi25 = phi25()
self.phi50 = phi50()
self.xRoot = xRoot()
self.zRoot = zRoot()
self.lHT = lHT()
self.cHT = cHT()
self.dihedral = dihedral()
self.LoD = LoD()
self.location = location()
# Aerodynamic
self.cLalpha = cLalpha()
self.cLMAX = cLMAX()
#self.cLMIN = cLMIN()
self.cLMAXht = cLMAXht()
self.CLalphaHTP_TO = CLalphaHTP_TO()
self.CLalphaHTP_CR = CLalphaHTP_CR()
self.CLalphaHTP_L = CLalphaHTP_L()
self.formFactor = formFactor()
self.reynoldsNr = reynoldsNr()
self.cfLAM = cfLAM()
self.cfTURB = cfTURB()
self.cD0c = cD0c()
self.nLam = nLam()
self.detadalpha = detadalpha()
self.DPR = DPR()
# Airfoil
self.airfoilr = airfoil(self) # Root Airfoil
self.airfoilt = airfoil(self) # Tip Airfoil
# Elevator
self.elevator = elevator(self)
def __init__(self, aircraft):
'''
will initialize instance mainly used for documentation
links to the aircraft instance
initiates the airfoil class
'''
component.__init__(self)
self.id = 'htp'
self.aircraft = aircraft
self.level = 2
self.UID = uID(cpacsPath='/cpacs/toolspecific/vampZero/htpUID')
# Mass
self.mHtp = mHtp(
cpacsPath=
'/cpacs/vehicles/aircraft/model/analyses/massBreakdown/mOEM/mEM/mStructure/mWingsStructure/mWingStructure[2]/massDescription['
+ self.id + '_mass]/mass')
self.parentUID = parameter(
value=self.id,
cpacsPath=
'/cpacs/vehicles/aircraft/model/analyses/massBreakdown/mOEM/mEM/mStructure/mWingsStructure/mWingStructure[2]/massDescription['
+ self.id + '_mass]/parentUID')
# CoG
self.posCoG = posCoG(
cpacsPath=
'/cpacs/vehicles/aircraft/model/analyses/massBreakdown/mOEM/mEM/mStructure/mWingsStructure/mWingStructure[2]/massDescription['
+ self.id + '_mass]/location/x')
# Geometry
self.refArea = refArea()
self.refAreaTakeOff = refAreaTakeOff()
self.refAreaCruise = refAreaCruise()
self.refAreaLanding = refAreaLanding()
self.expArea = expArea()
self.wetArea = wetArea()
self.aspectRatio = aspectRatio()
self.span = span()
self.taperRatio = taperRatio()
self.tcAVG = tcAVG()
self.cRoot = cRoot()
self.cTip = cTip()
self.cMAC = cMAC()
self.yMAC = yMAC()
self.xMAC = xMAC()
self.xMAC25 = xMAC25()
self.phiLE = phiLE()
self.phiTE = phiTE()
self.phi25 = phi25()
self.phi50 = phi50()
self.xRoot = xRoot()
self.zRoot = zRoot()
self.lHT = lHT()
self.cHT = cHT()
self.dihedral = dihedral()
self.LoD = LoD()
self.location = location()
# Aerodynamic
self.cLalpha = cLalpha()
self.cLMAX = cLMAX()
#self.cLMIN = cLMIN()
self.cLMAXht = cLMAXht()
self.CLalphaHTP_TO = CLalphaHTP_TO()
self.CLalphaHTP_CR = CLalphaHTP_CR()
self.CLalphaHTP_L = CLalphaHTP_L()
self.formFactor = formFactor()
self.reynoldsNr = reynoldsNr()
self.cfLAM = cfLAM()
self.cfTURB = cfTURB()
self.cD0c = cD0c()
self.nLam = nLam()
self.detadalpha = detadalpha()
self.DPR = DPR()
# Airfoil
self.airfoilr = airfoil(self) # Root Airfoil
self.airfoilt = airfoil(self) # Tip Airfoil
# Elevator
self.elevator = elevator(self)
def __init__(self, aircraft):
'''
will initialize instance mainly used for documentation
links to the aircraft instance
initiates the airfoil class
@Method: Component Constructor
'''
component.__init__(self)
self.id = 'tool'
self.aircraft = aircraft
self.level = 2
#===============================================================================
#Aircraft stuff
#===============================================================================
self.aircraftName = parameter(
cpacsPath='/cpacs/vehicles/aircraft/model/name', value='VAMPzero')
self.aircraftDES = parameter(
cpacsPath='/cpacs/vehicles/aircraft/model/description',
value='VAMPzero')
#commented out due to introduction of xMAC25
#self.refX = parameter(cpacsPath = '/cpacs/vehicles/aircraft/model/reference/point/x',value='0.0')
self.refY = parameter(
cpacsPath='/cpacs/vehicles/aircraft/model/reference/point/y',
value='0.0')
self.refZ = parameter(
cpacsPath='/cpacs/vehicles/aircraft/model/reference/point/z',
value='0.0')
#=======================================================================
# Lifting Line
#=======================================================================
self.liliname = parameter(
cpacsPath='/cpacs/toolspecific/liftingLine/tool/name',
value='LIFTING_LINE')
self.liliversion = parameter(
cpacsPath='/cpacs/toolspecific/liftingLine/tool/version',
value='2.3.1')
self.liliuID = parameter(
cpacsPath='/cpacs/toolspecific/liftingLine/aircraftModelUID',
value='A320modelID'
) #@todo: Model ID is not available at Time of Construction
self.liliSpaneling = parameter(
cpacsPath=
'/cpacs/toolspecific/liftingLine/toolParameters/wingPanelings/wingPaneling[1]/spanwise',
value='5')
self.liliCpaneling = parameter(
cpacsPath=
'/cpacs/toolspecific/liftingLine/toolParameters/wingPanelings/wingPaneling[1]/chordwise',
value='5')
self.lilimode = parameter(
cpacsPath=
'/cpacs/toolspecific/liftingLine/toolParameters/archiveMode',
value='1')
#self.lilipolars = parameter(cpacsPath = '/cpacs/toolspecific/liftingLine/performanceMaps',value='')
#=======================================================================
#Header Information
#=======================================================================
self.name = parameter(cpacsPath='/cpacs/header/name',
value='Export from')
self.creator = parameter(cpacsPath='/cpacs/header/creator',
value='VAMPzero')
self.timestamp = parameter(cpacsPath='/cpacs/header/timestamp',
value='2010-12-31T12:00:00')
self.version = parameter(cpacsPath='/cpacs/header/version',
value='0.1')
self.version = parameter(cpacsPath='/cpacs/header/cpacsVersion',
value='2.0')
def createFlapsSBW(parentWingCPACS, parentWingVAMPzero, myFlap):
'''
This is the main export method for the wings flaps for the strut braced wing.
The SBW does not feature any inner flaps as no kink exists
@todo: it is possible that the inner flap overlaps the kink area
'''
cpacsPath = '/cpacs/vehicles/aircraft/model/wings/wing[' + parentWingVAMPzero.id + ']'
cpacsWing = getObjfromXpath(parentWingCPACS, cpacsPath)
cpacsComponentSegment = cpacsWing.get_componentSegments().get_componentSegment()[0]
myFlaps = []
# Initialization, i.e. fetching values throughout the code
xsis = []
etas = []
for i in range(3,7):
try:
xsis.append(eval(cpacsComponentSegment.get_structure().get_spars().get_sparPositions().get_sparPosition()[i].get_xsi().valueOf_))
etas.append(eval(cpacsComponentSegment.get_structure().get_spars().get_sparPositions().get_sparPosition()[i].get_eta().valueOf_))
except IndexError:
pass
xsiSpar_interp = scipy.interpolate.interp1d(etas, xsis)
yFus = parentWingVAMPzero.yFuselage.getValue()
span = parentWingVAMPzero.span.getValue() / 2.
etaFus = yFus / span
etaKink = parentWingVAMPzero.etaKink.getValue()
cRoot = parentWingVAMPzero.cRoot.getValue()
cKink = calcChordLengthAtEta(etaKink, parentWingVAMPzero, cpacsWing)
innerFlapArea = myFlap.refAreaInnerFlap.getValue()
outerFlapArea = myFlap.refAreaOuterFlap.getValue()
# OuterFlap
sparOffset = 0.08
# maxX is the maximum extension of the flap
maxX = cRoot * (1. - (xsis[0] + sparOffset)) /2.
newOuterFlapArea = innerFlapArea + outerFlapArea
cRootOuterFlap = cRoot * (1. - (xsis[0] + sparOffset))
# Determine the total flap span by iteration
calcArea = 0.
spanOuterFlap = 0.
absSparOffset = sparOffset * cKink
# Obtain the maxEtaValue. This is forwarded from the ailerons export routine.
maxEta = myFlap.maxEta.getValue() - 0.02
# the tip root length is a function of the span (as the xsi location of the spar changes)
while abs(calcArea - newOuterFlapArea) > 0.1:
spanOuterFlap += .01
cTipOuterFlap = calcChordLengthAtEta(etaFus + spanOuterFlap / span, parentWingVAMPzero, cpacsWing) * (1 - xsiSpar_interp(etaFus + spanOuterFlap / span)) - absSparOffset
oldcalcArea = calcArea
calcArea = spanOuterFlap * (cRootOuterFlap + cTipOuterFlap) / 2.
if calcArea < oldcalcArea:
log.warning('VAMPzero FLAP: Outer Flap Area can not be established decreasing spar offset by 2% chord!')
sparOffset = sparOffset - 0.02
calcArea = 0.
spanOuterFlap = 0.
absSparOffset = sparOffset * cKink
break
if spanOuterFlap / span + etaFus > maxEta:
log.warning('VAMPzero FLAP: Outer Flap overlaps with the aileron decreasing spar offset by 2% chord!')
sparOffset = sparOffset - 0.02
calcArea = 0.
spanOuterFlap = 0.
absSparOffset = sparOffset * cRoot
break
while abs(calcArea - newOuterFlapArea) > 0.1:
spanOuterFlap += .01
cTipOuterFlap = calcChordLengthAtEta(etaFus + spanOuterFlap / span, parentWingVAMPzero, cpacsWing) * (1 - xsiSpar_interp(etaFus + spanOuterFlap / span)) - absSparOffset
oldcalcArea = calcArea
calcArea = spanOuterFlap * (cRootOuterFlap + cTipOuterFlap) / 2.
if calcArea < oldcalcArea:
log.warning('VAMPzero FLAP: Outer Flap Area can not be established! Continuing with outerArea = %s' % str(calcArea))
newOuterFlapArea = calcArea
break
if spanOuterFlap / span + etaFus + 0.06 > maxEta:
log.warning('VAMPzero FLAP: Outer Flap overlaps with the aileron! Continuing with outerArea = %s' % str(calcArea))
newOuterFlapArea = calcArea
break
# Determine the number of flaps
# The aspect ratio of a flap should not be higher than 6.
nOuterFlaps = spanOuterFlap ** 2. / (newOuterFlapArea * 6.)
log.debug('VAMPzero FLAP: Exporting %s Flaps outboard of the engine for an area of %s m2.' % (str(nOuterFlaps), str(newOuterFlapArea)))
# 1 Flap
if nOuterFlaps <= 1.:
# the inner border eta is located at the kink
innerEtaLE = etaFus + 0.05
innerXsiLE = xsis[0] + sparOffset
# the outer border eta is determined from the span of the outer flap
outerEtaLE = etaFus + 0.05 + spanOuterFlap / span
outerXsiLE = 1 - (cTipOuterFlap / calcChordLengthAtEta(outerEtaLE, parentWingVAMPzero, cpacsWing))
#Fowler Motion is restricted to 75% of the flap depth
innerX = (1. - innerXsiLE) * calcChordLengthAtEta(innerEtaLE, parentWingVAMPzero, cpacsWing) * 0.5
outerX = (1. - outerXsiLE) * calcChordLengthAtEta(outerEtaLE, parentWingVAMPzero, cpacsWing) * 0.5
myFlaps.append(createFlap('outerFlap1', parentWingVAMPzero.id, innerEtaLE, innerXsiLE, outerEtaLE, outerXsiLE, maxX, appendInnerCruiseRoller=True, type='flap', innerX=innerX, outerX=outerX))
# 2 Flaps
elif nOuterFlaps > 1. and nOuterFlaps <= 2.:
# the inner border eta is located at the kink
innerEtaLE = etaFus + 0.05
innerXsiLE = xsis[2] + sparOffset
# the outer border eta is determined from the half span of the total outer flap
outerEtaLE = etaFus + 0.05 + spanOuterFlap / (2*span)
outerXsiLE = xsiSpar_interp(outerEtaLE) + absSparOffset / calcChordLengthAtEta(outerEtaLE, parentWingVAMPzero, cpacsWing)
#Fowler Motion is restricted to 75% of the flap depth
innerX = (1. - innerXsiLE) * calcChordLengthAtEta(innerEtaLE, parentWingVAMPzero, cpacsWing) * 0.5
outerX = (1. - outerXsiLE) * calcChordLengthAtEta(outerEtaLE, parentWingVAMPzero, cpacsWing) * 0.5
myFlaps.append(createFlap('outerFlap1', parentWingVAMPzero.id, innerEtaLE, innerXsiLE, outerEtaLE, outerXsiLE, maxX, type='flap', innerX=innerX, outerX=outerX))
# new inner is the old outer
innerEtaLE = outerEtaLE
innerXsiLE = outerXsiLE
# the outer border eta is determined from the full span of the total outer flap
outerEtaLE = etaFus + 0.05 + spanOuterFlap / span
outerXsiLE = 1 - (cTipOuterFlap / calcChordLengthAtEta(outerEtaLE, parentWingVAMPzero, cpacsWing))
#Fowler Motion is restricted to 75% of the flap depth
innerX = (1. - innerXsiLE) * calcChordLengthAtEta(innerEtaLE, parentWingVAMPzero, cpacsWing) * 0.5
outerX = (1. - outerXsiLE) * calcChordLengthAtEta(outerEtaLE, parentWingVAMPzero, cpacsWing) * 0.5
myFlaps.append(createFlap('outerFlap2', parentWingVAMPzero.id, innerEtaLE, innerXsiLE, outerEtaLE, outerXsiLE, maxX, appendInnerCruiseRoller=True, type='flap', innerX=innerX, outerX=outerX))
# n Flaps
elif nOuterFlaps > 2. :
n = int(ceil(nOuterFlaps))
# First Flap
# the inner border eta is located at the kink
innerEtaLE = etaFus + 0.05
innerXsiLE = xsis[2] + sparOffset
# the outer border eta is determined from the half span of the total outer flap
outerEtaLE = etaFus + 0.05 + spanOuterFlap / (n*span)
outerXsiLE = xsiSpar_interp(outerEtaLE) + absSparOffset / calcChordLengthAtEta(outerEtaLE, parentWingVAMPzero, cpacsWing)
#Fowler Motion is restricted to 75% of the flap depth
innerX = (1. - innerXsiLE) * calcChordLengthAtEta(innerEtaLE, parentWingVAMPzero, cpacsWing) * 0.5
outerX = (1. - outerXsiLE) * calcChordLengthAtEta(outerEtaLE, parentWingVAMPzero, cpacsWing) * 0.5
myFlaps.append(createFlap('outerFlap1', parentWingVAMPzero.id, innerEtaLE, innerXsiLE, outerEtaLE, outerXsiLE, maxX, type='flap', innerX=innerX, outerX=outerX))
for i in range(2, n):
# nth Flap
# new inner is the old outer
innerEtaLE = outerEtaLE
innerXsiLE = outerXsiLE
# the outer border eta is determined from the full span of the total outer flap
outerEtaLE = innerEtaLE + spanOuterFlap / n / span
outerXsiLE = xsiSpar_interp(outerEtaLE) + absSparOffset / calcChordLengthAtEta(outerEtaLE, parentWingVAMPzero, cpacsWing)
#Fowler Motion is restricted to 75% of the flap depth
innerX = (1. - innerXsiLE) * calcChordLengthAtEta(innerEtaLE, parentWingVAMPzero, cpacsWing) * 0.5
outerX = (1. - outerXsiLE) * calcChordLengthAtEta(outerEtaLE, parentWingVAMPzero, cpacsWing) * 0.5
myFlaps.append(createFlap('outerFlap' + str(i), parentWingVAMPzero.id, innerEtaLE, innerXsiLE, outerEtaLE, outerXsiLE, maxX, appendInnerCruiseRoller=True, type='flap', innerX=innerX, outerX=outerX))
# Last Flap
# new inner is the old outer
innerEtaLE = outerEtaLE
innerXsiLE = outerXsiLE
# the outer border eta is determined from the full span of the total outer flap
outerEtaLE = etaFus + 0.05 + spanOuterFlap / span
outerXsiLE = 1 - (cTipOuterFlap / calcChordLengthAtEta(outerEtaLE, parentWingVAMPzero, cpacsWing))
#Fowler Motion is restricted to 75% of the flap depth
innerX = (1. - innerXsiLE) * calcChordLengthAtEta(innerEtaLE, parentWingVAMPzero, cpacsWing) * 0.5
outerX = (1. - outerXsiLE) * calcChordLengthAtEta(outerEtaLE, parentWingVAMPzero, cpacsWing) * 0.5
myFlaps.append(createFlap('outerFlap' + str(n), parentWingVAMPzero.id, innerEtaLE, innerXsiLE, outerEtaLE, outerXsiLE, maxX, appendInnerCruiseRoller=True, type='flap', innerX=innerX, outerX=outerX))
# Output to Spoiler
# as the spoiler is relying on data of the flaps some basic information is written
# back to the the VAMPzero components
parentWingVAMPzero.spoiler.outerEta = parameter(parent=parentWingVAMPzero.spoiler, value=outerEtaLE, unit='', status='calc', doc='The outermost eta coordinate of all flaps. This overlaps with the outer eta coordinate of the spoiler')
spoilerChord = maxX
parentWingVAMPzero.spoiler.chord = parameter(spoilerChord, 'm', 'calc', 'The absolute chord of the spoiler: 5% of the kink chord length + 50% of the innerFlap Chord length', parent=parentWingVAMPzero.spoiler)
# Output to CPACS
if type(cpacsComponentSegment.get_controlSurfaces()) == NoneType:
cpacsComponentSegment.set_controlSurfaces(controlSurfacesType())
if type(cpacsComponentSegment.get_controlSurfaces().get_trailingEdgeDevices()) == NoneType:
cpacsComponentSegment.get_controlSurfaces().set_trailingEdgeDevices(trailingEdgeDevicesType())
log.debug('VAMPzero SLAT: Exporting %s Flaps to CPACS.' % (str(len(myFlaps))))
for flap in myFlaps:
cpacsComponentSegment.get_controlSurfaces().get_trailingEdgeDevices().add_trailingEdgeDevice(flap)
def __init__(self, aircraft):
'''
will initialize instance mainly used for documentation
links to the aircraft instance
initiates the airfoil class
@Method: Component Constructor
'''
component.__init__(self)
self.id = 'strut'
self.aircraft = aircraft
self.level = 2
# Mass
self.mStrut = mStrut(
cpacsPath=
'/cpacs/vehicles/aircraft/model/analyses/massBreakdown/mOEM/mEM/mStructure/mWingsStructure/mWingStructure[4]/massDescription['
+ self.id + '_mass]/mass')
self.parentUID = parameter(
value=self.id,
cpacsPath=
'/cpacs/vehicles/aircraft/model/analyses/massBreakdown/mOEM/mEM/mStructure/mWingsStructure/mWingStructure[4]/massDescription['
+ self.id + '_mass]/parentUID')
# Geometry
self.refArea = refArea()
self.expArea = expArea()
self.wetArea = wetArea()
self.aspectRatio = aspectRatio()
self.span = span()
self.depth = depth()
self.active = active()
self.taperRatio = taperRatio()
self.tcAVG = tcAVG()
self.cRoot = cRoot()
self.cTip = cTip()
self.cMAC = cMAC()
self.yMAC = yMAC()
self.xMAC = xMAC()
self.xMAC25 = xMAC25()
self.phiLE = phiLE()
self.phiTE = phiTE()
self.phi25 = phi25()
self.phi50 = phi50()
self.dihedral = dihedral()
self.twist = twist()
self.etaStrut = etaStrut()
self.xRoot = xRoot()
self.yRoot = yRoot()
self.zRoot = zRoot()
self.xTip = xTip()
self.yTip = yTip()
self.zTip = zTip()
# Aerodynamics
self.formFactor = formFactor()
self.reynoldsNr = reynoldsNr()
self.cfLAM = cfLAM()
self.cfTURB = cfTURB()
self.cD0c = cD0c()
self.nLam = nLam()
# Airfoil
self.airfoilr = airfoil(self, position='root')
self.airfoilt = airfoil(self, position='tip')
def __init__(self, aircraft):
'''
will initialize instance mainly used for documentation
links to the aircraft instance
initiates the airfoil class
@Method: Component Constructor
'''
component.__init__(self)
self.id = 'wing'
self.aircraft = aircraft
self.level = 2
self.UID = uID(cpacsPath='/cpacs/toolspecific/vampZero/wingUID')
# Mass
self.mWing = mWing(
cpacsPath=
'/cpacs/vehicles/aircraft/model/analyses/massBreakdown/mOEM/mEM/mStructure/mWingsStructure/mWingStructure[1]/massDescription['
+ self.id + '_mass]/mass')
self.parentUID = parameter(
value=self.id,
cpacsPath=
'/cpacs/vehicles/aircraft/model/analyses/massBreakdown/mOEM/mEM/mStructure/mWingsStructure/mWingStructure[1]/massDescription['
+ self.id + '_mass]/parentUID')
# CoG
self.posCoG = posCoG(
cpacsPath=
'/cpacs/vehicles/aircraft/model/analyses/massBreakdown/mOEM/mEM/mStructure/mWingsStructure/mWingStructure[1]/massDescription['
+ self.id + '_mass]/location/x')
# Geometry
self.location = location()
self.refArea = refArea(
cpacsPath='/cpacs/vehicles/aircraft/model/reference/area')
self.expArea = expArea()
self.wetArea = wetArea()
self.aspectRatio = aspectRatio()
self.span = span()
self.taperRatio = taperRatio()
self.tcAVG = tcAVG()
self.cRoot = cRoot()
self.cTip = cTip()
self.cMAC = cMAC(
cpacsPath='/cpacs/vehicles/aircraft/model/reference/length')
self.yMAC = yMAC()
self.xMAC = xMAC()
self.xMAC25 = xMAC25(
cpacsPath='/cpacs/vehicles/aircraft/model/reference/point/x')
self.phiLE = phiLE()
self.phiTE = phiTE()
self.phi25 = phi25()
self.phi50 = phi50()
self.dihedral = dihedral()
self.twist = twist()
self.xRoot = xRoot()
self.zRoot = zRoot()
self.LoD = LoD()
self.etaKink = etaKink()
self.etaEngine = etaEngine()
# Aerodynamics
self.cLalpha = cLalpha()
self.cLMAX = cLMAX()
self.formFactor = formFactor()
self.reynoldsNr = reynoldsNr(
) # cpacsPath='/cpacs/vehicles/aircraft/model/global/aeroPerformanceMap/reynoldsNumber')
self.cfLAM = cfLAM()
self.cfTURB = cfTURB()
self.cD0c = cD0c()
self.oswald = oswald()
self.cDMINoffset = cDMINoffset()
self.nLam = nLam()
self.cDw = cDw()
self.machDD = machDD()
self.machCrit = machCrit()
self.cM0CR = cM0CR()
self.cM0TO = cM0TO()
self.cM0L = cM0L()
# Airfoil
self.airfoilr = airfoil(self, position='root')
self.airfoilt = airfoil(self, position='tip')
# Aileron
self.aileron = aileron(self)
# Slat
self.slat = slat(self)
# Flap
self.flap = flap(self)
# Spoiler
self.spoiler = spoiler(self)
# CPACS Stuff for Export to higher Level
self.xFuselage = xFuselage()
self.yFuselage = yFuselage()
self.zFuselage = zFuselage()
self.cFuselage = cFuselage()
self.xKink = xKink()
self.yKink = yKink()
self.zKink = zKink()
self.cKink = cKink()
self.xTip = xTip()
self.zTip = zTip()
def createAileron(parentWingCPACS, parentWingVAMPzero, myAileron):
'''
This is the main export method for the wings aileron
'''
cpacsPath = '/cpacs/vehicles/aircraft/model/wings/wing[' + parentWingVAMPzero.id + ']'
cpacsWing = getObjfromXpath(parentWingCPACS, cpacsPath)
cpacsComponentSegment = cpacsWing.get_componentSegments().get_componentSegment()[0]
# Header
myName = stringBaseType(None, None, None, 'aileron')
myDescription = stringBaseType(None, None, None, 'aileron from VAMPzero')
myParentUID = stringUIDBaseType(None, None, 'True', None, 'wing_Cseg')
# Initialization, i.e. fetching values throughout the code
xsis = []
etas = []
for i in range(3,7):
try:
xsis.append(eval(cpacsComponentSegment.get_structure().get_spars().get_sparPositions().get_sparPosition()[i].get_xsi().valueOf_))
etas.append(eval(cpacsComponentSegment.get_structure().get_spars().get_sparPositions().get_sparPosition()[i].get_eta().valueOf_))
except IndexError:
pass
xsiSpar_interp = scipy.interpolate.interp1d(etas, xsis)
sparOffset = 0.1
wingSpan = parentWingVAMPzero.span.getValue() / 2.
# Outer Shape
# The outer border eta station is set to 96 percent
outerEtaLE = 0.96
# The outer chord station is determined from the wing's chord at eta = 0.96
# and the rear spar location + the spar offset
outerXsiLE = xsiSpar_interp(0.96) + sparOffset
outerWingChord = calcChordLengthAtEta(outerEtaLE, parentWingVAMPzero, cpacsWing)
cTip = (1 - outerXsiLE) * outerWingChord
# now we need to determine the necessary span for the aileron by gently increasing the span
# this is an iterative process as the chord of the aileron is a function of the inbound span
aileronArea = parentWingVAMPzero.aileron.refArea.getValue()
delta = 0.01
calcArea = 0.
while abs(calcArea - aileronArea) > 0.1:
if delta > outerEtaLE:
parentWingVAMPzero.log.warning('VAMPzero EXPORT: Cannot determine the span of the aileron')
parentWingVAMPzero.log.warning('VAMPzero EXPORT: aileronArea= '+str(aileronArea))
parentWingVAMPzero.log.warning('VAMPzero EXPORT: Decreasing Spar Offset')
sparOffset = sparOffset - 0.02
delta = 0.01
innerEtaLE = outerEtaLE - delta
innerXsiLE = xsiSpar_interp(innerEtaLE) + sparOffset
innerWingChord = calcChordLengthAtEta(innerEtaLE, parentWingVAMPzero, cpacsWing)
cRoot = (1 - innerXsiLE) * innerWingChord
calcArea = (cTip + cRoot) / 2 * (outerEtaLE - innerEtaLE) * wingSpan
delta += 0.005
# start outer shape
myleadingEdgeShape = leadingEdgeShapeType(relHeightLE=doubleBaseType(valueOf_=str(0.5)), xsiUpperSkin=doubleBaseType(valueOf_=str(0.85)), xsiLowerSkin=doubleBaseType(valueOf_=str(0.85)))
innerBorder = controlSurfaceBorderTrailingEdgeType(etaLE=doubleBaseType(valueOf_=str(innerEtaLE)), etaTE=doubleBaseType(valueOf_=str(innerEtaLE)), xsiLE=doubleBaseType(valueOf_=str(innerXsiLE)), leadingEdgeShape=myleadingEdgeShape)
outerBorder = controlSurfaceBorderTrailingEdgeType(etaLE=doubleBaseType(valueOf_=str(outerEtaLE)), etaTE=doubleBaseType(valueOf_=str(outerEtaLE)), xsiLE=doubleBaseType(valueOf_=str(outerXsiLE)), leadingEdgeShape=myleadingEdgeShape)
myOuterShape = controlSurfaceOuterShapeTrailingEdgeType(innerBorder=innerBorder, outerBorder=outerBorder)
# structure
myStructure = wingComponentSegmentStructureType()
cpacsAileron = trailingEdgeDeviceType(uID='aileronUID', name=myName, description=myDescription, parentUID=myParentUID, outerShape=myOuterShape, structure=myStructure)
createAileronStructure(cpacsAileron)
# Forward information about innerEtaLE to the flap
parentWingVAMPzero.flap.maxEta = parameter(value=innerEtaLE, doc='This it the inner position of the aileron, the flap may not exceed it')
# moveables
deltaEta = outerEtaLE - innerEtaLE
innerParentXsi = xsiSpar_interp(innerEtaLE + 0.3 * deltaEta) + 0.02
outerParentXsi = xsiSpar_interp(innerEtaLE + 0.7 * deltaEta) + 0.02
createPath(cpacsAileron, 'aileron')
createTracks(cpacsAileron, 'aileron')
createActuators(cpacsAileron, 'aileron', [innerParentXsi, outerParentXsi])
if type(cpacsComponentSegment.get_controlSurfaces()) == NoneType:
cpacsComponentSegment.set_controlSurfaces(controlSurfacesType())
if type(cpacsComponentSegment.get_controlSurfaces().get_trailingEdgeDevices()) == NoneType:
cpacsComponentSegment.get_controlSurfaces().set_trailingEdgeDevices(trailingEdgeDevicesType())
cpacsComponentSegment.get_controlSurfaces().get_trailingEdgeDevices().add_trailingEdgeDevice(cpacsAileron)
def createAileron(parentWingCPACS, parentWingVAMPzero, myAileron):
'''
This is the main export method for the wings aileron
'''
cpacsPath = '/cpacs/vehicles/aircraft/model/wings/wing[' + parentWingVAMPzero.id + ']'
cpacsWing = getObjfromXpath(parentWingCPACS, cpacsPath)
cpacsComponentSegment = cpacsWing.get_componentSegments(
).get_componentSegment()[0]
#===========================================================================
# Header
#===========================================================================
myName = stringBaseType(None, None, None, 'aileron')
myDescription = stringBaseType(None, None, None, 'aileron from VAMPzero')
myParentUID = stringUIDBaseType(None, None, 'True', None, 'wing_Cseg')
#===========================================================================
# Initialization, i.e. fetching values throughout the code
#===========================================================================
xsiSparRoot = eval(cpacsComponentSegment.get_structure().get_spars(
).get_sparPositions().get_sparPosition()[3].get_xsi().valueOf_)
xsiSparFuselage = eval(cpacsComponentSegment.get_structure().get_spars(
).get_sparPositions().get_sparPosition()[4].get_xsi().valueOf_)
xsiSparKink = eval(cpacsComponentSegment.get_structure().get_spars(
).get_sparPositions().get_sparPosition()[5].get_xsi().valueOf_)
xsiSparTip = eval(cpacsComponentSegment.get_structure().get_spars(
).get_sparPositions().get_sparPosition()[6].get_xsi().valueOf_)
etaSparRoot = eval(cpacsComponentSegment.get_structure().get_spars(
).get_sparPositions().get_sparPosition()[3].get_eta().valueOf_)
etaSparFuselage = eval(cpacsComponentSegment.get_structure().get_spars(
).get_sparPositions().get_sparPosition()[4].get_eta().valueOf_)
etaSparKink = eval(cpacsComponentSegment.get_structure().get_spars(
).get_sparPositions().get_sparPosition()[5].get_eta().valueOf_)
etaSparTip = eval(cpacsComponentSegment.get_structure().get_spars(
).get_sparPositions().get_sparPosition()[6].get_eta().valueOf_)
xsiSpar_interp = scipy.interpolate.interp1d(
[etaSparRoot, etaSparFuselage, etaSparKink, etaSparTip],
[xsiSparRoot, xsiSparFuselage, xsiSparKink, xsiSparTip])
sparOffset = 0.1
wingSpan = parentWingVAMPzero.span.getValue() / 2.
#===========================================================================
# Outer Shape
#===========================================================================
# The outer border eta station is set to 96 percent
outerEtaLE = 0.96
# The outer chord station is determined from the wing's chord at eta = 0.96
# and the rear spar location + the spar offset
outerXsiLE = xsiSpar_interp(0.96) + sparOffset
outerWingChord = calcChordLengthAtEta(outerEtaLE, parentWingVAMPzero,
cpacsWing)
cTip = (1 - outerXsiLE) * outerWingChord
# now we need to determine the necessary span for the aileron by gently increasing the span
# this is an iterative process as the chord of the aileron is a function of the inbound span
aileronArea = parentWingVAMPzero.aileron.refArea.getValue()
delta = 0.01
calcArea = 0.
while abs(calcArea - aileronArea) > 0.1:
if delta > outerEtaLE:
parentWingVAMPzero.log.warning(
'VAMPzero EXPORT: Cannot determine the span of the aileron')
parentWingVAMPzero.log.warning('VAMPzero EXPORT: aileronArea= ' +
str(aileronArea))
parentWingVAMPzero.log.warning(
'VAMPzero EXPORT: Decreasing Spar Offset')
sparOffset = sparOffset - 0.02
delta = 0.01
innerEtaLE = outerEtaLE - delta
innerXsiLE = xsiSpar_interp(innerEtaLE) + sparOffset
innerWingChord = calcChordLengthAtEta(innerEtaLE, parentWingVAMPzero,
cpacsWing)
cRoot = (1 - innerXsiLE) * innerWingChord
calcArea = (cTip + cRoot) / 2 * (outerEtaLE - innerEtaLE) * wingSpan
delta += 0.005
# start outer shape
myleadingEdgeShape = leadingEdgeShapeType(
relHeightLE=doubleBaseType(valueOf_=str(0.5)),
xsiUpperSkin=doubleBaseType(valueOf_=str(0.85)),
xsiLowerSkin=doubleBaseType(valueOf_=str(0.85)))
innerBorder = controlSurfaceBorderTrailingEdgeType(
etaLE=doubleBaseType(valueOf_=str(innerEtaLE)),
etaTE=doubleBaseType(valueOf_=str(innerEtaLE)),
xsiLE=doubleBaseType(valueOf_=str(innerXsiLE)),
leadingEdgeShape=myleadingEdgeShape)
outerBorder = controlSurfaceBorderTrailingEdgeType(
etaLE=doubleBaseType(valueOf_=str(outerEtaLE)),
etaTE=doubleBaseType(valueOf_=str(outerEtaLE)),
xsiLE=doubleBaseType(valueOf_=str(outerXsiLE)),
leadingEdgeShape=myleadingEdgeShape)
myOuterShape = controlSurfaceOuterShapeTrailingEdgeType(
innerBorder=innerBorder, outerBorder=outerBorder)
# structure
myStructure = wingComponentSegmentStructureType()
cpacsAileron = trailingEdgeDeviceType(uID='aileronUID',
name=myName,
description=myDescription,
parentUID=myParentUID,
outerShape=myOuterShape,
structure=myStructure)
createAileronStructure(cpacsAileron)
# Forward information about innerEtaLE to the flap
parentWingVAMPzero.flap.maxEta = parameter(
value=innerEtaLE,
doc=
'This it the inner position of the aileron, the flap may not exceed it'
)
# moveables
deltaEta = outerEtaLE - innerEtaLE
innerParentXsi = xsiSpar_interp(innerEtaLE + 0.3 * deltaEta) + 0.02
outerParentXsi = xsiSpar_interp(innerEtaLE + 0.7 * deltaEta) + 0.02
createPath(cpacsAileron, 'aileron')
createTracks(cpacsAileron, 'aileron')
createActuators(cpacsAileron, 'aileron', [innerParentXsi, outerParentXsi])
if type(cpacsComponentSegment.get_controlSurfaces()) == NoneType:
cpacsComponentSegment.set_controlSurfaces(controlSurfacesType())
if type(cpacsComponentSegment.get_controlSurfaces().
get_trailingEdgeDevices()) == NoneType:
cpacsComponentSegment.get_controlSurfaces().set_trailingEdgeDevices(
trailingEdgeDevicesType())
cpacsComponentSegment.get_controlSurfaces().get_trailingEdgeDevices(
).add_trailingEdgeDevice(cpacsAileron)
def __init__(self, aircraft):
"""
will initialize instance mainly used for documentation
links to the aircraft instance
initiates the airfoil class
@Method: Component Constructor
"""
component.__init__(self)
self.id = "strut"
self.aircraft = aircraft
self.level = 2
# Mass
self.mStrut = mStrut(
cpacsPath="/cpacs/vehicles/aircraft/model/analyses/massBreakdown/mOEM/mEM/mStructure/mWingsStructure/mWingStructure[4]/massDescription["
+ self.id
+ "_mass]/mass"
)
self.parentUID = parameter(
value=self.id,
cpacsPath="/cpacs/vehicles/aircraft/model/analyses/massBreakdown/mOEM/mEM/mStructure/mWingsStructure/mWingStructure[4]/massDescription["
+ self.id
+ "_mass]/parentUID",
)
# Geometry
self.refArea = refArea()
self.expArea = expArea()
self.wetArea = wetArea()
self.aspectRatio = aspectRatio()
self.span = span()
self.depth = depth()
self.active = active()
self.taperRatio = taperRatio()
self.tcAVG = tcAVG()
self.cRoot = cRoot()
self.cTip = cTip()
self.cMAC = cMAC()
self.yMAC = yMAC()
self.xMAC = xMAC()
self.xMAC25 = xMAC25()
self.phiLE = phiLE()
self.phiTE = phiTE()
self.phi25 = phi25()
self.phi50 = phi50()
self.dihedral = dihedral()
self.twist = twist()
self.etaStrut = etaStrut()
self.xRoot = xRoot()
self.yRoot = yRoot()
self.zRoot = zRoot()
self.xTip = xTip()
self.yTip = yTip()
self.zTip = zTip()
# Aerodynamics
self.formFactor = formFactor()
self.reynoldsNr = reynoldsNr()
self.cfLAM = cfLAM()
self.cfTURB = cfTURB()
self.cD0c = cD0c()
self.nLam = nLam()
# Airfoil
self.airfoilr = airfoil(self, position="root")
self.airfoilt = airfoil(self, position="tip")
def __init__(self, aircraft):
'''
will initialize instance mainly used for documentation
links to the aircraft instance
initiates the airfoil class
@Method: Component Constructor
'''
component.__init__(self)
self.id = 'vtp'
self.aircraft = aircraft
self.level = 2
self.UID = uID(cpacsPath='/cpacs/toolspecific/vampZero/vtpUID')
# Mass
self.mVtp = mVtp(
cpacsPath='/cpacs/vehicles/aircraft/model/analyses/massBreakdown/mOEM/mEM/mStructure/mWingsStructure/mWingStructure[3]/massDescription['+self.id+'_mass]/mass')
self.parentUID = parameter(value=self.id,
cpacsPath='/cpacs/vehicles/aircraft/model/analyses/massBreakdown/mOEM/mEM/mStructure/mWingsStructure/mWingStructure[3]/massDescription['+self.id+'_mass]/parentUID')
# CoG
self.posCoG = posCoG(
cpacsPath='/cpacs/vehicles/aircraft/model/analyses/massBreakdown/mOEM/mEM/mStructure/mWingsStructure/mWingStructure[3]/massDescription['+self.id+'_mass]/location/x')
# Geometry
self.refArea = refArea()
self.refAreaTakeOff = refAreaTakeOff()
self.refAreaLanding = refAreaLanding()
self.expArea = expArea()
self.wetArea = wetArea()
self.aspectRatio = aspectRatio()
self.span = span()
self.taperRatio = taperRatio()
self.tcAVG = tcAVG()
self.cRoot = cRoot()
self.cTip = cTip()
self.cMAC = cMAC()
self.yMAC = yMAC()
self.xMAC = xMAC()
self.phiLE = phiLE()
self.phiTE = phiTE()
self.phi25 = phi25()
self.phi50 = phi50()
self.xRoot = xRoot()
self.zRoot = zRoot()
self.lVT = lVT()
self.cVT = cVT()
self.LoD = LoD()
# Aerodynamics
self.cLalpha = cLalpha()
self.cLMAX = cLMAX()
self.formFactor = formFactor()
self.reynoldsNr = reynoldsNr()
self.cfLAM = cfLAM()
self.cfTURB = cfTURB()
self.cD0c = cD0c()
self.nLam = nLam()
# Airfoil
self.airfoilr = airfoil(self) # Root Airfoil
self.airfoilt = airfoil(self) # Tip Airfoil
# Rudder
self.rudder = rudder(self)
def createFlapsAdvDoubleTrapezoid(parentWingCPACS, parentWingVAMPzero, myFlap):
'''
This is the main export method for the wings flaps
@todo: it is possible that the inner flap overlaps the kink area
'''
cpacsPath = '/cpacs/vehicles/aircraft/model/wings/wing[' + parentWingVAMPzero.id + ']'
cpacsWing = getObjfromXpath(parentWingCPACS, cpacsPath)
cpacsComponentSegment = cpacsWing.get_componentSegments().get_componentSegment()[0]
myFlaps = []
# Initialization, i.e. fetching values throughout the code
xsiSparRoot = eval(cpacsComponentSegment.get_structure().get_spars().get_sparPositions().get_sparPosition()[3].get_xsi().valueOf_)
xsiSparFuselage = eval(cpacsComponentSegment.get_structure().get_spars().get_sparPositions().get_sparPosition()[4].get_xsi().valueOf_)
xsiSparKink = eval(cpacsComponentSegment.get_structure().get_spars().get_sparPositions().get_sparPosition()[5].get_xsi().valueOf_)
xsiSparTip = eval(cpacsComponentSegment.get_structure().get_spars().get_sparPositions().get_sparPosition()[6].get_xsi().valueOf_)
etaSparRoot = eval(cpacsComponentSegment.get_structure().get_spars().get_sparPositions().get_sparPosition()[3].get_eta().valueOf_)
etaSparFuselage = eval(cpacsComponentSegment.get_structure().get_spars().get_sparPositions().get_sparPosition()[4].get_eta().valueOf_)
etaSparKink = eval(cpacsComponentSegment.get_structure().get_spars().get_sparPositions().get_sparPosition()[5].get_eta().valueOf_)
etaSparTip = eval(cpacsComponentSegment.get_structure().get_spars().get_sparPositions().get_sparPosition()[6].get_eta().valueOf_)
xsiSpar_interp = scipy.interpolate.interp1d([etaSparRoot, etaSparFuselage, etaSparKink, etaSparTip], [xsiSparRoot, xsiSparFuselage, xsiSparKink, xsiSparTip])
yFus = parentWingVAMPzero.yFuselage.getValue()
span = parentWingVAMPzero.span.getValue() / 2.
etaFus = yFus / span
etaKink = parentWingVAMPzero.etaKink.getValue()
cRoot = parentWingVAMPzero.cRoot.getValue()
cKink = calcChordLengthAtEta(etaKink, parentWingVAMPzero, cpacsWing)
phiLE = parentWingVAMPzero.phiLE.getValue()
innerFlapArea = myFlap.refAreaInnerFlap.getValue()
outerFlapArea = myFlap.refAreaOuterFlap.getValue()
# InnerFlap
# Determine the number of flaps
# The aspect ratio of a flap should not be higher than 9.
sparOffset = 0.08
if phiLE > 0.:
# if the wing is backward swept the chord of the inner flap equals the chord at kink behind the rear spar + sparOffset
cInnerFlap = cKink * (1. - (xsiSparKink + sparOffset))
elif phiLE < 0.:
# if the wing is forward swept the chord of the inner flap is similar to the chord of the root section behind the spar + 20%
cInnerFlap = cRoot * (1. - (xsiSparFuselage + 0.2))
# maxX is the maximum extension of the flap
maxX = cInnerFlap / 2.
# The area of the innerFlap is determined by the chord and span
# The difference in the area will be substituted by the outerFlap Area
spanInnerFlap = (etaKink - etaFus) * span
newInnerFlapArea = cInnerFlap * spanInnerFlap
deltaArea = innerFlapArea - newInnerFlapArea
nInnerFlaps = spanInnerFlap ** 2. / (newInnerFlapArea * 9.)
log.debug('VAMPzero FLAP: Exporting %s Flaps inside of the engine for an area of %s m2.' % (str(nInnerFlaps), str(innerFlapArea)))
if nInnerFlaps < 1.:
# the inner border eta is located at the intersection with the fuselage
innerEtaLE = etaFus
# the inner border xsi is determined from the wing chord at the fuselage station (which equals CRoot)
innerXsiLE = 1. - (cInnerFlap / cRoot)
# the outer border eta is determined from the span of the inner flap
outerEtaLE = etaKink
# the outer xsi location is found from the avg chord and the chord length at that station
outerXsiLE = 1. - (cInnerFlap / cKink)
myFlaps.append(createFlap('innerFlap', parentWingVAMPzero.id, innerEtaLE, innerXsiLE, outerEtaLE, outerXsiLE, maxX, type='innerFlap', innerX=maxX, outerX=maxX))
elif nInnerFlaps > 1. and nInnerFlaps <= 2.:
# First Flap
# the inner border eta is located at the intersection with the fuselage
innerEtaLE = etaFus
# the inner border xsi is determined from the wing chord at the fuselage station (which equals CRoot)
innerXsiLE = 1. - (cInnerFlap / cRoot)
# the outer border eta is determined from the span of the inner flap
outerEtaLE = etaFus + (etaKink - etaFus) / 2.
# the outer xsi location is found from the avg chord and the chord length at that station
outerXsiLE = 1. - (cInnerFlap / calcChordLengthAtEta(outerEtaLE, parentWingVAMPzero, cpacsWing))
myFlaps.append(createFlap('innerFlap1', parentWingVAMPzero.id, innerEtaLE, innerXsiLE, outerEtaLE, outerXsiLE, maxX, type='innerFlap', innerX=maxX, outerX=maxX))
# Second Flap
# new inner is old outer
innerEtaLE = outerEtaLE
# new outer is kink
outerEtaLE = etaKink
# new inner is old outer
innerXsiLE = outerXsiLE
# new outer is kink
outerXsiLE = 1. - (cInnerFlap / cKink)
myFlaps.append(createFlap('innerFlap2', parentWingVAMPzero.id, innerEtaLE, innerXsiLE, outerEtaLE, outerXsiLE, maxX, type='flap', innerX=maxX, outerX=maxX))
# OuterFlap
newOuterFlapArea = outerFlapArea + deltaArea
cRootOuterFlap = cKink * (1. - (xsiSparKink + sparOffset))
# Determine the total flap span by iteration
calcArea = 0.
spanOuterFlap = 0.
absSparOffset = sparOffset * cKink
# Obtain the maxEtaValue. This is forwarded from the ailerons export routine.
maxEta = myFlap.maxEta.getValue() - 0.02
# the tip root length is a function of the span (as the xsi location of the spar changes)
while abs(calcArea - newOuterFlapArea) > 0.1:
spanOuterFlap += .01
cTipOuterFlap = calcChordLengthAtEta(etaKink + spanOuterFlap / span, parentWingVAMPzero, cpacsWing) * (1 - xsiSpar_interp(etaKink + spanOuterFlap / span)) - absSparOffset
oldcalcArea = calcArea
calcArea = spanOuterFlap * (cRootOuterFlap + cTipOuterFlap) / 2.
if calcArea < oldcalcArea:
log.warning('VAMPzero FLAP: Outer Flap Area can not be established decreasing spar offset by 2% chord!')
sparOffset = sparOffset - 0.02
calcArea = 0.
spanOuterFlap = 0.
absSparOffset = sparOffset * cKink
break
if spanOuterFlap / span + etaKink > maxEta:
log.warning('VAMPzero FLAP: Outer Flap overlaps with the aileron decreasing spar offset by 2% chord!')
sparOffset = sparOffset - 0.02
calcArea = 0.
spanOuterFlap = 0.
absSparOffset = sparOffset * cKink
break
while abs(calcArea - newOuterFlapArea) > 0.1:
spanOuterFlap += .01
cTipOuterFlap = calcChordLengthAtEta(etaKink + spanOuterFlap / span, parentWingVAMPzero, cpacsWing) * (1 - xsiSpar_interp(etaKink + spanOuterFlap / span)) - absSparOffset
oldcalcArea = calcArea
calcArea = spanOuterFlap * (cRootOuterFlap + cTipOuterFlap) / 2.
if calcArea < oldcalcArea:
log.warning('VAMPzero FLAP: Outer Flap Area can not be established! Continuing with outerArea = %s' % str(calcArea))
newOuterFlapArea = calcArea
break
if spanOuterFlap / span + etaKink > maxEta:
log.warning('VAMPzero FLAP: Outer Flap overlaps with the aileron! Continuing with outerArea = %s' % str(calcArea))
newOuterFlapArea = calcArea
break
# Determine the number of flaps
# The aspect ratio of a flap should not be higher than 9.
nOuterFlaps = spanOuterFlap ** 2. / (newOuterFlapArea * 9.)
log.debug('VAMPzero FLAP: Exporting %s Flaps outboard of the engine for an area of %s m2.' % (str(nOuterFlaps), str(newOuterFlapArea)))
# 1 Flap
if nOuterFlaps <= 1.:
# the inner border eta is located at the kink
innerEtaLE = etaKink
innerXsiLE = xsiSparKink + sparOffset
# the outer border eta is determined from the span of the outer flap
outerEtaLE = etaKink + spanOuterFlap / span
outerXsiLE = 1 - (cTipOuterFlap / calcChordLengthAtEta(outerEtaLE, parentWingVAMPzero, cpacsWing))
#Fowler Motion is restricted to 75% of the flap depth
innerX = (1. - innerXsiLE) * calcChordLengthAtEta(innerEtaLE, parentWingVAMPzero, cpacsWing) * 0.5
outerX = (1. - outerXsiLE) * calcChordLengthAtEta(outerEtaLE, parentWingVAMPzero, cpacsWing) * 0.5
myFlaps.append(createFlap('outerFlap1', parentWingVAMPzero.id, innerEtaLE, innerXsiLE, outerEtaLE, outerXsiLE, maxX, appendInnerCruiseRoller=True, type='flap', innerX=innerX, outerX=outerX))
# 2 Flaps
elif nOuterFlaps > 1. and nOuterFlaps <= 2.:
# the inner border eta is located at the kink
innerEtaLE = etaKink
innerXsiLE = xsiSparKink + sparOffset
# the outer border eta is determined from the half span of the total outer flap
outerEtaLE = etaKink + spanOuterFlap / (2. * span)
outerXsiLE = xsiSpar_interp(outerEtaLE) + absSparOffset / calcChordLengthAtEta(outerEtaLE, parentWingVAMPzero, cpacsWing)
#Fowler Motion is restricted to 75% of the flap depth
innerX = (1. - innerXsiLE) * calcChordLengthAtEta(innerEtaLE, parentWingVAMPzero, cpacsWing) * 0.5
outerX = (1. - outerXsiLE) * calcChordLengthAtEta(outerEtaLE, parentWingVAMPzero, cpacsWing) * 0.5
myFlaps.append(createFlap('outerFlap1', parentWingVAMPzero.id, innerEtaLE, innerXsiLE, outerEtaLE, outerXsiLE, maxX, type='flap', innerX=innerX, outerX=outerX))
# new inner is the old outer
innerEtaLE = outerEtaLE
innerXsiLE = outerXsiLE
# the outer border eta is determined from the full span of the total outer flap
outerEtaLE = etaKink + spanOuterFlap / span
outerXsiLE = 1 - (cTipOuterFlap / calcChordLengthAtEta(outerEtaLE, parentWingVAMPzero, cpacsWing))
#Fowler Motion is restricted to 75% of the flap depth
innerX = (1. - innerXsiLE) * calcChordLengthAtEta(innerEtaLE, parentWingVAMPzero, cpacsWing) * 0.5
outerX = (1. - outerXsiLE) * calcChordLengthAtEta(outerEtaLE, parentWingVAMPzero, cpacsWing) * 0.5
myFlaps.append(createFlap('outerFlap2', parentWingVAMPzero.id, innerEtaLE, innerXsiLE, outerEtaLE, outerXsiLE, maxX, appendInnerCruiseRoller=True, type='flap', innerX=innerX, outerX=outerX))
# n Flaps
elif nOuterFlaps > 2. :
n = int(ceil(nOuterFlaps))
# First Flap
# the inner border eta is located at the kink
innerEtaLE = etaKink
innerXsiLE = xsiSparKink + sparOffset
# the outer border eta is determined from the half span of the total outer flap
outerEtaLE = etaKink + spanOuterFlap / n / span
outerXsiLE = xsiSpar_interp(outerEtaLE) + absSparOffset / calcChordLengthAtEta(outerEtaLE, parentWingVAMPzero, cpacsWing)
#Fowler Motion is restricted to 75% of the flap depth
innerX = (1. - innerXsiLE) * calcChordLengthAtEta(innerEtaLE, parentWingVAMPzero, cpacsWing) * 0.5
outerX = (1. - outerXsiLE) * calcChordLengthAtEta(outerEtaLE, parentWingVAMPzero, cpacsWing) * 0.5
myFlaps.append(createFlap('outerFlap1', parentWingVAMPzero.id, innerEtaLE, innerXsiLE, outerEtaLE, outerXsiLE, maxX, type='flap', innerX=innerX, outerX=outerX))
for i in range(2, n):
# nth Flap
# new inner is the old outer
innerEtaLE = outerEtaLE
innerXsiLE = outerXsiLE
# the outer border eta is determined from the full span of the total outer flap
outerEtaLE = innerEtaLE + spanOuterFlap / n / span
outerXsiLE = xsiSpar_interp(outerEtaLE) + absSparOffset / calcChordLengthAtEta(outerEtaLE, parentWingVAMPzero, cpacsWing)
#Fowler Motion is restricted to 75% of the flap depth
innerX = (1. - innerXsiLE) * calcChordLengthAtEta(innerEtaLE, parentWingVAMPzero, cpacsWing) * 0.5
outerX = (1. - outerXsiLE) * calcChordLengthAtEta(outerEtaLE, parentWingVAMPzero, cpacsWing) * 0.5
myFlaps.append(createFlap('outerFlap' + str(i), parentWingVAMPzero.id, innerEtaLE, innerXsiLE, outerEtaLE, outerXsiLE, maxX, appendInnerCruiseRoller=True, type='flap', innerX=innerX, outerX=outerX))
# Last Flap
# new inner is the old outer
innerEtaLE = outerEtaLE
innerXsiLE = outerXsiLE
# the outer border eta is determined from the full span of the total outer flap
outerEtaLE = etaKink + spanOuterFlap / span
outerXsiLE = 1 - (cTipOuterFlap / calcChordLengthAtEta(outerEtaLE, parentWingVAMPzero, cpacsWing))
#Fowler Motion is restricted to 75% of the flap depth
innerX = (1. - innerXsiLE) * calcChordLengthAtEta(innerEtaLE, parentWingVAMPzero, cpacsWing) * 0.5
outerX = (1. - outerXsiLE) * calcChordLengthAtEta(outerEtaLE, parentWingVAMPzero, cpacsWing) * 0.5
myFlaps.append(createFlap('outerFlap' + str(n), parentWingVAMPzero.id, innerEtaLE, innerXsiLE, outerEtaLE, outerXsiLE, maxX, appendInnerCruiseRoller=True, type='flap', innerX=innerX, outerX=outerX))
# Output to Spoiler
# as the spoiler is relying on data of the flaps some basic information is written
# back to the the VAMPzero components
parentWingVAMPzero.spoiler.outerEta = parameter(outerEtaLE, '', 'calc', 'The outermost eta coordinate of all flaps. This overlaps with the outer eta coordinate of the spoiler', parent=parentWingVAMPzero.spoiler)
spoilerChord = 0.05 * cKink + 0.5 * cInnerFlap
parentWingVAMPzero.spoiler.chord = parameter(spoilerChord, 'm', 'calc', 'The absolute chord of the spoiler: 5% of the kink chord length + 50% of the innerFlap Chord length', parent=parentWingVAMPzero.spoiler)
# Output to CPACS
if type(cpacsComponentSegment.get_controlSurfaces()) == NoneType:
cpacsComponentSegment.set_controlSurfaces(controlSurfacesType())
if type(cpacsComponentSegment.get_controlSurfaces().get_trailingEdgeDevices()) == NoneType:
cpacsComponentSegment.get_controlSurfaces().set_trailingEdgeDevices(trailingEdgeDevicesType())
log.debug('VAMPzero SLAT: Exporting %s Flaps to CPACS.' % (str(len(myFlaps))))
for flap in myFlaps:
cpacsComponentSegment.get_controlSurfaces().get_trailingEdgeDevices().add_trailingEdgeDevice(flap)
def __init__(self, aircraft):
'''
will initialize instance mainly used for documentation
links to the aircraft instance
initiates the airfoil class
@Method: Component Constructor
'''
component.__init__(self)
self.id = 'wing'
self.aircraft = aircraft
self.level = 2
self.UID = uID(cpacsPath='/cpacs/toolspecific/vampZero/wingUID')
# Mass
self.mWing = mWing(
cpacsPath='/cpacs/vehicles/aircraft/model/analyses/massBreakdown/mOEM/mEM/mStructure/mWingsStructure/mWingStructure[1]/massDescription['+self.id+'_mass]/mass')
self.parentUID = parameter(value=self.id,
cpacsPath='/cpacs/vehicles/aircraft/model/analyses/massBreakdown/mOEM/mEM/mStructure/mWingsStructure/mWingStructure[1]/massDescription['+self.id+'_mass]/parentUID')
# CoG
self.posCoG = posCoG(
cpacsPath='/cpacs/vehicles/aircraft/model/analyses/massBreakdown/mOEM/mEM/mStructure/mWingsStructure/mWingStructure[1]/massDescription['+self.id+'_mass]/location/x')
# Geometry
self.location = location()
self.refArea = refArea(cpacsPath='/cpacs/vehicles/aircraft/model/reference/area')
self.expArea = expArea()
self.wetArea = wetArea()
self.aspectRatio = aspectRatio()
self.span = span()
self.taperRatio = taperRatio()
self.tcAVG = tcAVG()
self.cRoot = cRoot()
self.cTip = cTip()
self.cMAC = cMAC(cpacsPath='/cpacs/vehicles/aircraft/model/reference/length')
self.yMAC = yMAC()
self.xMAC = xMAC()
self.xMAC25 = xMAC25(cpacsPath='/cpacs/vehicles/aircraft/model/reference/point/x')
self.phiLE = phiLE()
self.phiTE = phiTE()
self.phi25 = phi25()
self.phi50 = phi50()
self.dihedral = dihedral()
self.twist = twist()
self.xRoot = xRoot()
self.zRoot = zRoot()
self.LoD = LoD()
self.etaKink = etaKink()
self.etaEngine = etaEngine()
# Aerodynamics
self.cLalpha = cLalpha()
self.cLMAX = cLMAX()
self.formFactor = formFactor()
self.reynoldsNr = reynoldsNr() # cpacsPath='/cpacs/vehicles/aircraft/model/global/aeroPerformanceMap/reynoldsNumber')
self.cfLAM = cfLAM()
self.cfTURB = cfTURB()
self.cD0c = cD0c()
self.oswald = oswald()
self.cDMINoffset = cDMINoffset()
self.nLam = nLam()
self.cDw = cDw()
self.machDD = machDD()
self.machCrit = machCrit()
self.cM0CR = cM0CR()
self.cM0TO = cM0TO()
self.cM0L = cM0L()
# Airfoil
self.airfoilr = airfoil(self, position='root')
self.airfoilt = airfoil(self, position='tip')
# Aileron
self.aileron = aileron(self)
# Slat
self.slat = slat(self)
# Flap
self.flap = flap(self)
# Spoiler
self.spoiler = spoiler(self)
# CPACS Stuff for Export to higher Level
self.xFuselage = xFuselage()
self.yFuselage = yFuselage()
self.zFuselage = zFuselage()
self.cFuselage = cFuselage()
self.xKink = xKink()
self.yKink = yKink()
self.zKink = zKink()
self.cKink = cKink()
self.xTip = xTip()
self.zTip = zTip()
def createFlapsSBW(parentWingCPACS, parentWingVAMPzero, myFlap):
'''
This is the main export method for the wings flaps
@todo: it is possible that the inner flap overlaps the kink area
'''
cpacsPath = '/cpacs/vehicles/aircraft/model/wings/wing[' + parentWingVAMPzero.id + ']'
cpacsWing = getObjfromXpath(parentWingCPACS, cpacsPath)
cpacsComponentSegment = cpacsWing.get_componentSegments(
).get_componentSegment()[0]
myFlaps = []
#===========================================================================
# Initialization, i.e. fetching values throughout the code
#===========================================================================
xsiSparRoot = eval(cpacsComponentSegment.get_structure().get_spars(
).get_sparPositions().get_sparPosition()[3].get_xsi().valueOf_)
xsiSparFuselage = eval(cpacsComponentSegment.get_structure().get_spars(
).get_sparPositions().get_sparPosition()[4].get_xsi().valueOf_)
xsiSparKink = eval(cpacsComponentSegment.get_structure().get_spars(
).get_sparPositions().get_sparPosition()[5].get_xsi().valueOf_)
xsiSparTip = eval(cpacsComponentSegment.get_structure().get_spars(
).get_sparPositions().get_sparPosition()[6].get_xsi().valueOf_)
etaSparRoot = eval(cpacsComponentSegment.get_structure().get_spars(
).get_sparPositions().get_sparPosition()[3].get_eta().valueOf_)
etaSparFuselage = eval(cpacsComponentSegment.get_structure().get_spars(
).get_sparPositions().get_sparPosition()[4].get_eta().valueOf_)
etaSparKink = eval(cpacsComponentSegment.get_structure().get_spars(
).get_sparPositions().get_sparPosition()[5].get_eta().valueOf_)
etaSparTip = eval(cpacsComponentSegment.get_structure().get_spars(
).get_sparPositions().get_sparPosition()[6].get_eta().valueOf_)
xsiSpar_interp = scipy.interpolate.interp1d(
[etaSparRoot, etaSparFuselage, etaSparKink, etaSparTip],
[xsiSparRoot, xsiSparFuselage, xsiSparKink, xsiSparTip])
yFus = parentWingVAMPzero.yFuselage.getValue()
span = parentWingVAMPzero.span.getValue() / 2.
etaFus = yFus / span
etaKink = parentWingVAMPzero.etaKink.getValue()
cRoot = parentWingVAMPzero.cRoot.getValue()
cKink = calcChordLengthAtEta(etaKink, parentWingVAMPzero, cpacsWing)
phiLE = parentWingVAMPzero.phiLE.getValue()
innerFlapArea = myFlap.refAreaInnerFlap.getValue()
outerFlapArea = myFlap.refAreaOuterFlap.getValue()
#===========================================================================
# OuterFlap
#===========================================================================
sparOffset = 0.08
# maxX is the maximum extension of the flap
maxX = cRoot * (1. - (xsiSparRoot + sparOffset)) / 2.
newOuterFlapArea = innerFlapArea + outerFlapArea
cRootOuterFlap = cRoot * (1. - (xsiSparRoot + sparOffset))
#===========================================================================
# Determine the total flap span by iteration
#===========================================================================
calcArea = 0.
spanOuterFlap = 0.
absSparOffset = sparOffset * cKink
# Obtain the maxEtaValue. This is forwarded from the ailerons export routine.
maxEta = myFlap.maxEta.getValue() - 0.02
# the tip root length is a function of the span (as the xsi location of the spar changes)
while abs(calcArea - newOuterFlapArea) > 0.1:
spanOuterFlap += .01
cTipOuterFlap = calcChordLengthAtEta(
etaFus + spanOuterFlap / span, parentWingVAMPzero, cpacsWing
) * (1 - xsiSpar_interp(etaFus + spanOuterFlap / span)) - absSparOffset
oldcalcArea = calcArea
calcArea = spanOuterFlap * (cRootOuterFlap + cTipOuterFlap) / 2.
if calcArea < oldcalcArea:
log.warning(
'VAMPzero FLAP: Outer Flap Area can not be established decreasing spar offset by 2% chord!'
)
sparOffset = sparOffset - 0.02
calcArea = 0.
spanOuterFlap = 0.
absSparOffset = sparOffset * cKink
break
if spanOuterFlap / span + etaFus > maxEta:
log.warning(
'VAMPzero FLAP: Outer Flap overlaps with the aileron decreasing spar offset by 2% chord!'
)
sparOffset = sparOffset - 0.02
calcArea = 0.
spanOuterFlap = 0.
absSparOffset = sparOffset * cRoot
break
while abs(calcArea - newOuterFlapArea) > 0.1:
spanOuterFlap += .01
cTipOuterFlap = calcChordLengthAtEta(
etaKink + spanOuterFlap / span, parentWingVAMPzero, cpacsWing) * (
1 -
xsiSpar_interp(etaKink + spanOuterFlap / span)) - absSparOffset
oldcalcArea = calcArea
calcArea = spanOuterFlap * (cRootOuterFlap + cTipOuterFlap) / 2.
if calcArea < oldcalcArea:
log.warning(
'VAMPzero FLAP: Outer Flap Area can not be established! Continuing with outerArea = %s'
% str(calcArea))
newOuterFlapArea = calcArea
break
if spanOuterFlap / span + etaFus > maxEta:
log.warning(
'VAMPzero FLAP: Outer Flap overlaps with the aileron! Continuing with outerArea = %s'
% str(calcArea))
newOuterFlapArea = calcArea
break
#===========================================================================
# Determine the number of flaps
# The aspect ratio of a flap should not be higher than 6.
#===========================================================================
nOuterFlaps = spanOuterFlap**2. / (newOuterFlapArea * 6.)
log.debug(
'VAMPzero FLAP: Exporting %s Flaps outboard of the engine for an area of %s m2.'
% (str(nOuterFlaps), str(newOuterFlapArea)))
#===========================================================================
# 1 Flap
#===========================================================================
if nOuterFlaps <= 1.:
# the inner border eta is located at the kink
innerEtaLE = etaFus + 0.05
innerXsiLE = xsiSparRoot + sparOffset
# the outer border eta is determined from the span of the outer flap
outerEtaLE = etaFus + 0.05 + spanOuterFlap / span
outerXsiLE = 1 - (cTipOuterFlap / calcChordLengthAtEta(
outerEtaLE, parentWingVAMPzero, cpacsWing))
#Fowler Motion is restricted to 75% of the flap depth
innerX = (1. - innerXsiLE) * calcChordLengthAtEta(
innerEtaLE, parentWingVAMPzero, cpacsWing) * 0.5
outerX = (1. - outerXsiLE) * calcChordLengthAtEta(
outerEtaLE, parentWingVAMPzero, cpacsWing) * 0.5
myFlaps.append(
createFlap('outerFlap1',
parentWingVAMPzero.id,
innerEtaLE,
innerXsiLE,
outerEtaLE,
outerXsiLE,
maxX,
appendInnerCruiseRoller=True,
type='flap',
innerX=innerX,
outerX=outerX))
#===========================================================================
# 2 Flaps
#===========================================================================
elif nOuterFlaps > 1. and nOuterFlaps <= 2.:
# the inner border eta is located at the kink
innerEtaLE = etaFus + 0.05
innerXsiLE = xsiSparKink + sparOffset
# the outer border eta is determined from the half span of the total outer flap
outerEtaLE = etaFus + 0.05 + spanOuterFlap / (2 * span)
outerXsiLE = xsiSpar_interp(
outerEtaLE) + absSparOffset / calcChordLengthAtEta(
outerEtaLE, parentWingVAMPzero, cpacsWing)
#Fowler Motion is restricted to 75% of the flap depth
innerX = (1. - innerXsiLE) * calcChordLengthAtEta(
innerEtaLE, parentWingVAMPzero, cpacsWing) * 0.5
outerX = (1. - outerXsiLE) * calcChordLengthAtEta(
outerEtaLE, parentWingVAMPzero, cpacsWing) * 0.5
myFlaps.append(
createFlap('outerFlap1',
parentWingVAMPzero.id,
innerEtaLE,
innerXsiLE,
outerEtaLE,
outerXsiLE,
maxX,
type='flap',
innerX=innerX,
outerX=outerX))
# new inner is the old outer
innerEtaLE = outerEtaLE
innerXsiLE = outerXsiLE
# the outer border eta is determined from the full span of the total outer flap
outerEtaLE = etaFus + 0.05 + spanOuterFlap / span
outerXsiLE = 1 - (cTipOuterFlap / calcChordLengthAtEta(
outerEtaLE, parentWingVAMPzero, cpacsWing))
#Fowler Motion is restricted to 75% of the flap depth
innerX = (1. - innerXsiLE) * calcChordLengthAtEta(
innerEtaLE, parentWingVAMPzero, cpacsWing) * 0.5
outerX = (1. - outerXsiLE) * calcChordLengthAtEta(
outerEtaLE, parentWingVAMPzero, cpacsWing) * 0.5
myFlaps.append(
createFlap('outerFlap2',
parentWingVAMPzero.id,
innerEtaLE,
innerXsiLE,
outerEtaLE,
outerXsiLE,
maxX,
appendInnerCruiseRoller=True,
type='flap',
innerX=innerX,
outerX=outerX))
#===========================================================================
# n Flaps
#===========================================================================
elif nOuterFlaps > 2.:
n = int(ceil(nOuterFlaps))
# First Flap
# the inner border eta is located at the kink
innerEtaLE = etaFus + 0.05
innerXsiLE = xsiSparKink + sparOffset
# the outer border eta is determined from the half span of the total outer flap
outerEtaLE = etaFus + 0.05 + spanOuterFlap / (n * span)
outerXsiLE = xsiSpar_interp(
outerEtaLE) + absSparOffset / calcChordLengthAtEta(
outerEtaLE, parentWingVAMPzero, cpacsWing)
#Fowler Motion is restricted to 75% of the flap depth
innerX = (1. - innerXsiLE) * calcChordLengthAtEta(
innerEtaLE, parentWingVAMPzero, cpacsWing) * 0.5
outerX = (1. - outerXsiLE) * calcChordLengthAtEta(
outerEtaLE, parentWingVAMPzero, cpacsWing) * 0.5
myFlaps.append(
createFlap('outerFlap1',
parentWingVAMPzero.id,
innerEtaLE,
innerXsiLE,
outerEtaLE,
outerXsiLE,
maxX,
type='flap',
innerX=innerX,
outerX=outerX))
for i in range(2, n):
# nth Flap
# new inner is the old outer
innerEtaLE = outerEtaLE
innerXsiLE = outerXsiLE
# the outer border eta is determined from the full span of the total outer flap
outerEtaLE = innerEtaLE + spanOuterFlap / n / span
outerXsiLE = xsiSpar_interp(
outerEtaLE) + absSparOffset / calcChordLengthAtEta(
outerEtaLE, parentWingVAMPzero, cpacsWing)
#Fowler Motion is restricted to 75% of the flap depth
innerX = (1. - innerXsiLE) * calcChordLengthAtEta(
innerEtaLE, parentWingVAMPzero, cpacsWing) * 0.5
outerX = (1. - outerXsiLE) * calcChordLengthAtEta(
outerEtaLE, parentWingVAMPzero, cpacsWing) * 0.5
myFlaps.append(
createFlap('outerFlap' + str(i),
parentWingVAMPzero.id,
innerEtaLE,
innerXsiLE,
outerEtaLE,
outerXsiLE,
maxX,
appendInnerCruiseRoller=True,
type='flap',
innerX=innerX,
outerX=outerX))
# Last Flap
# new inner is the old outer
innerEtaLE = outerEtaLE
innerXsiLE = outerXsiLE
# the outer border eta is determined from the full span of the total outer flap
outerEtaLE = etaFus + 0.05 + spanOuterFlap / span
outerXsiLE = 1 - (cTipOuterFlap / calcChordLengthAtEta(
outerEtaLE, parentWingVAMPzero, cpacsWing))
#Fowler Motion is restricted to 75% of the flap depth
innerX = (1. - innerXsiLE) * calcChordLengthAtEta(
innerEtaLE, parentWingVAMPzero, cpacsWing) * 0.5
outerX = (1. - outerXsiLE) * calcChordLengthAtEta(
outerEtaLE, parentWingVAMPzero, cpacsWing) * 0.5
myFlaps.append(
createFlap('outerFlap' + str(n),
parentWingVAMPzero.id,
innerEtaLE,
innerXsiLE,
outerEtaLE,
outerXsiLE,
maxX,
appendInnerCruiseRoller=True,
type='flap',
innerX=innerX,
outerX=outerX))
#===============================================================================
# Output to Spoiler
# as the spoiler is relying on data of the flaps some basic information is written
# back to the the VAMPzero components
#===============================================================================
parentWingVAMPzero.spoiler.outerEta = parameter(
parent=parentWingVAMPzero.spoiler,
value=outerEtaLE,
unit='',
status='calc',
doc=
'The outermost eta coordinate of all flaps. This overlaps with the outer eta coordinate of the spoiler'
)
spoilerChord = maxX
parentWingVAMPzero.spoiler.chord = parameter(
spoilerChord,
'm',
'calc',
'The absolute chord of the spoiler: 5% of the kink chord length + 50% of the innerFlap Chord length',
parent=parentWingVAMPzero.spoiler)
#===========================================================================
# Output to CPACS
#===========================================================================
if type(cpacsComponentSegment.get_controlSurfaces()) == NoneType:
cpacsComponentSegment.set_controlSurfaces(controlSurfacesType())
if type(cpacsComponentSegment.get_controlSurfaces().
get_trailingEdgeDevices()) == NoneType:
cpacsComponentSegment.get_controlSurfaces().set_trailingEdgeDevices(
trailingEdgeDevicesType())
log.debug('VAMPzero SLAT: Exporting %s Flaps to CPACS.' %
(str(len(myFlaps))))
for flap in myFlaps:
cpacsComponentSegment.get_controlSurfaces().get_trailingEdgeDevices(
).add_trailingEdgeDevice(flap)
def __init__(self, aircraft):
'''
will initialize instance mainly used for documentation
links to the aircraft instance
initiates the airfoil class
@Method: Component Constructor
'''
component.__init__(self)
self.id = 'tool'
self.aircraft = aircraft
self.level = 2
#===============================================================================
#Aircraft stuff
#===============================================================================
self.aircraftName = parameter(cpacsPath='/cpacs/vehicles/aircraft/model/name', value='VAMPzero')
self.aircraftDES = parameter(cpacsPath='/cpacs/vehicles/aircraft/model/description', value='VAMPzero')
#commented out due to introduction of xMAC25
#self.refX = parameter(cpacsPath = '/cpacs/vehicles/aircraft/model/reference/point/x',value='0.0')
self.refY = parameter(cpacsPath='/cpacs/vehicles/aircraft/model/reference/point/y', value='0.0')
self.refZ = parameter(cpacsPath='/cpacs/vehicles/aircraft/model/reference/point/z', value='0.0')
#=======================================================================
# Lifting Line
#=======================================================================
self.liliname = parameter(cpacsPath='/cpacs/toolspecific/liftingLine/tool/name', value='LIFTING_LINE')
self.liliversion = parameter(cpacsPath='/cpacs/toolspecific/liftingLine/tool/version', value='2.3.1')
self.liliuID = parameter(cpacsPath='/cpacs/toolspecific/liftingLine/aircraftModelUID',
value='A320modelID')#@todo: Model ID is not available at Time of Construction
self.liliSpaneling = parameter(
cpacsPath='/cpacs/toolspecific/liftingLine/toolParameters/wingPanelings/wingPaneling[1]/spanwise',
value='5')
self.liliCpaneling = parameter(
cpacsPath='/cpacs/toolspecific/liftingLine/toolParameters/wingPanelings/wingPaneling[1]/chordwise',
value='5')
self.lilimode = parameter(cpacsPath='/cpacs/toolspecific/liftingLine/toolParameters/archiveMode', value='1')
#self.lilipolars = parameter(cpacsPath = '/cpacs/toolspecific/liftingLine/performanceMaps',value='')
#=======================================================================
#Header Information
#=======================================================================
self.name = parameter(cpacsPath='/cpacs/header/name', value='Export from')
self.creator = parameter(cpacsPath='/cpacs/header/creator', value='VAMPzero')
self.timestamp = parameter(cpacsPath='/cpacs/header/timestamp', value='2010-12-31T12:00:00')
self.version = parameter(cpacsPath='/cpacs/header/version', value='0.1')
self.version = parameter(cpacsPath='/cpacs/header/cpacsVersion', value='2.0')
