def tas2ssec(tas, alt, oat, ias, speed_units=default_speed_units, alt_units=default_alt_units, temp_units=default_temp_units, press_units = default_press_units):
"""
Return static source position error as speed error, pressure error and altitude error at sea level.
Returns delta_Vpc, delta_Ps, delta_Hpc and Vc
delta_Vpc = error in airspeed = calibrated airspeed - indicated airspeed corrected for instrument error
delta_Ps = error in the pressure sensed by the static system = pressure sensed in the static system - ambient pressure
delta_Hpc = altitude error at sea level = actual altitude - altitude sensed by the static system
Vc = calibrated airspeed at the test point
"""
P0 = U.press_conv(constants.P0, from_units = 'pa', to_units = press_units)
cas = A.tas2cas(tas, alt, oat, speed_units=speed_units, alt_units=alt_units, temp_units=temp_units)
delta_Vpc = cas - ias
qcic = A.cas2dp(ias, speed_units=speed_units, press_units =press_units)
qc = A.cas2dp(cas, speed_units=speed_units, press_units = press_units)
delta_Ps = qc - qcic
delta_Hpc = SA.press2alt(P0 - delta_Ps, press_units = press_units)
# print 'Vc = %.1f, delta_vpc = %.1f, delta_ps = %.1f, delta_hpc = %.0f' % (cas, delta_Vpc, delta_Ps, delta_Hpc)
return delta_Vpc, delta_Ps, delta_Hpc, cas
def WOT_speed(altitude, weight = 2100, rpm = 2700, temp = 'std', \
temp_units = 'C', rv = 'F1', wing_area = 102, speed_units = 'kt', \
prop_eff = 0.78, MP_loss = 1.322, ram = 0.5, rated_power=275):
"""
Returns the predicted speed at full throttle.
The MP_loss is the MP lost in the induction tract at full throttle at 2700
rpm at MSL. The default value is from the Lycoming power charts.
The ram is the percentage of available ram recovery pressure that is
achieved in the MP.
"""
press = SA.alt2press(altitude, press_units='in HG')
MP_loss = MP_loss * (rpm / 2700.)**2
cas_guess = 300
error = 1
while error > 0.0001:
dp = A.cas2dp(cas_guess, press_units='in HG')
# print 'dp =', dp
ram_press = ram * dp
# print 'Ram rise =', ram_press
MP = press - MP_loss + ram_press
pwr = O.pwr(rpm, MP, altitude, temp=temp,
temp_units=temp_units) * rated_power / 180
print('MP =', MP, 'Power =', pwr)
tas = speed(altitude, weight, pwr, rpm, temp = temp, \
temp_units = temp_units, rv = rv, wing_area = wing_area, \
speed_units = speed_units, prop_eff = prop_eff)
cas = A.tas2cas(tas, altitude, temp=temp, temp_units=temp_units)
error = M.fabs((cas - cas_guess) / cas)
cas_guess = cas
# print 'CAS =', cas, 'TAS =', tas
return tas
def WOT_speed(altitude, weight = 2100, rpm = 2700, temp = 'std', \
temp_units = 'C', rv = 'F1', wing_area = 102, speed_units = 'kt', \
prop_eff = 0.78, MP_loss = 1.322, ram = 0.5, rated_power=275):
"""
Returns the predicted speed at full throttle.
The MP_loss is the MP lost in the induction tract at full throttle at 2700
rpm at MSL. The default value is from the Lycoming power charts.
The ram is the percentage of available ram recovery pressure that is
achieved in the MP.
"""
press = SA.alt2press(altitude, press_units = 'in HG')
MP_loss = MP_loss * (rpm / 2700.) ** 2
cas_guess = 300
error = 1
while error > 0.0001:
dp = A.cas2dp(cas_guess, press_units = 'in HG')
# print 'dp =', dp
ram_press = ram * dp
# print 'Ram rise =', ram_press
MP = press - MP_loss + ram_press
pwr = O.pwr(rpm, MP, altitude, temp = temp, temp_units = temp_units) * rated_power/180
print 'MP =', MP, 'Power =', pwr
tas = speed(altitude, weight, pwr, rpm, temp = temp, \
temp_units = temp_units, rv = rv, wing_area = wing_area, \
speed_units = speed_units, prop_eff = prop_eff)
cas = A.tas2cas(tas, altitude, temp = temp, temp_units = temp_units)
error = M.fabs((cas - cas_guess) / cas)
cas_guess = cas
# print 'CAS =', cas, 'TAS =', tas
return tas
def test_16(self):
# 1000 kt to pa
# truth value from NASA RP 1046
Value = A.cas2dp(1000, press_units='pa')
Truth = 249053
self.failUnless(RE(Value, Truth) <= 2e-5)
def test_15(self):
# 1000 kt to mm HG
# truth value from NASA RP 1046
Value = A.cas2dp(1000, press_units='mm HG')
Truth = 1868.05
self.failUnless(RE(Value, Truth) <= 2e-5)
def test_14(self):
# 1700 km/h to pa
# truth value from NASA RP 1046
Value = A.cas2dp(1700, press_units='pa', speed_units='km/h')
Truth = 203298
self.failUnless(RE(Value, Truth) <= 2e-5)
def test_13(self):
# 1700 km/h to mm HG
# truth value from NASA RP 1046
Value = A.cas2dp(1700, press_units='mm HG', speed_units='km/h')
Truth = 1524.86
self.failUnless(RE(Value, Truth) <= 2e-5)
def test_12(self):
# 1000 kt to psf
# truth value from NASA RP 1046
Value = A.cas2dp(1000, press_units='psf')
Truth = 5201.59
self.failUnless(RE(Value, Truth) <= 2e-5)
def test_05(self):
# 300 km/h to mm HG
# truth value from NASA RP 1046
Value = A.cas2dp(300, press_units='mm HG', speed_units='km/h')
Truth = 32.385
self.failUnless(RE(Value, Truth) <= 1e-5)
def test_03(self):
# 300 mph to in HG
# truth value from NASA RP 1046
Value = A.cas2dp(300, press_units='in HG', speed_units='mph')
Truth = 3.38145
self.failUnless(RE(Value, Truth) <= 1e-5)
def test_09(self):
# 1000 mph to in HG
# truth value from NASA RP 1046
Value = A.cas2dp(1000, press_units='in HG', speed_units='mph')
Truth = 52.5970
self.failUnless(RE(Value, Truth) <= 2e-5)
def test_08(self):
# 244 kt to pa
# truth value from NASA RP 1046
Value = A.cas2dp(244, press_units='pa')
Truth = 9983.7
self.failUnless(RE(Value, Truth) <= 2e-5)
def test_07(self):
# 280 kt to mm HG
# truth value from NASA RP 1046
Value = A.cas2dp(280, press_units='mm HG')
Truth = 99.671
self.assertTrue(RE(Value, Truth) <= 2e-5)
def test_01(self):
# 100 kt to lb/ft**3
# truth value from NASA RP 1046
Value = A.cas2dp(100, press_units='psf')
Truth = 34.0493
self.failUnless(RE(Value, Truth) <= 1e-5)
def test_10(self):
# 1000 mph to psf
# truth value from NASA RP 1046
Value = A.cas2dp(1000, press_units='psf', speed_units='mph')
Truth = 3719.98
self.failUnless(RE(Value, Truth) <= 2e-5)
def test_02(self):
# 600 kt to lb/ft**3
# truth value from NASA RP 1046
Value = A.cas2dp(700, press_units='psf', speed_units='mph')
Truth = 1540.37
self.failUnless(RE(Value, Truth) <= 1e-5)
def test_11(self):
# 1000 kt to in HG
# truth value from NASA RP 1046
Value = A.cas2dp(1000, press_units='in HG')
Truth = 73.5454
self.failUnless(RE(Value, Truth) <= 2e-5)
def test_06(self):
# 299 km/h to pa
# truth value from NASA RP 1046
Value = A.cas2dp(299, press_units='pa', speed_units='km/h')
Truth = 4288.5
self.assertTrue(RE(Value, Truth) <= 1e-5)
def test_04(self):
# 300 mph to in HG
# truth value from NASA RP 1046
Value = A.cas2dp(55, press_units='in HG', speed_units='kt')
Truth = .145052
self.failUnless(RE(Value, Truth) <= 1e-5)
def test_05(self):
# 300 km/h to mm HG
# truth value from NASA RP 1046
Value = A.cas2dp(300, press_units='mm HG', speed_units='km/h')
Truth = 32.385
self.assertTrue(RE(Value, Truth) <= 1e-5)
def test_06(self):
# 299 km/h to pa
# truth value from NASA RP 1046
Value = A.cas2dp(299, press_units='pa', speed_units='km/h')
Truth = 4288.5
self.failUnless(RE(Value, Truth) <= 1e-5)
def test_03(self):
# 300 mph to in HG
# truth value from NASA RP 1046
Value = A.cas2dp(300, press_units='in HG', speed_units='mph')
Truth = 3.38145
self.assertTrue(RE(Value, Truth) <= 1e-5)
def test_04(self):
# 300 mph to in HG
# truth value from NASA RP 1046
Value = A.cas2dp(55, press_units='in HG', speed_units='kt')
Truth = .145052
self.assertTrue(RE(Value, Truth) <= 1e-5)
def test_01(self):
# 100 kt to lb/ft**3
# truth value from NASA RP 1046
Value = A.cas2dp(100, press_units='psf')
Truth = 34.0493
self.assertTrue(RE(Value, Truth) <= 1e-5)
def test_02(self):
# 600 kt to lb/ft**3
# truth value from NASA RP 1046
Value = A.cas2dp(700, press_units='psf', speed_units='mph')
Truth = 1540.37
self.assertTrue(RE(Value, Truth) <= 1e-5)
def test_15(self):
# 1000 kt to mm HG
# truth value from NASA RP 1046
Value = A.cas2dp(1000, press_units='mm HG')
Truth = 1868.05
self.assertTrue(RE(Value, Truth) <= 2e-5)
def test_16(self):
# 1000 kt to pa
# truth value from NASA RP 1046
Value = A.cas2dp(1000, press_units='pa')
Truth = 249053
self.assertTrue(RE(Value, Truth) <= 2e-5)
def test_13(self):
# 1700 km/h to mm HG
# truth value from NASA RP 1046
Value = A.cas2dp(1700, press_units='mm HG', speed_units='km/h')
Truth = 1524.86
self.assertTrue(RE(Value, Truth) <= 2e-5)
def test_14(self):
# 1700 km/h to pa
# truth value from NASA RP 1046
Value = A.cas2dp(1700, press_units='pa', speed_units='km/h')
Truth = 203298
self.assertTrue(RE(Value, Truth) <= 2e-5)
def test_11(self):
# 1000 kt to in HG
# truth value from NASA RP 1046
Value = A.cas2dp(1000, press_units='in HG')
Truth = 73.5454
self.assertTrue(RE(Value, Truth) <= 2e-5)
def test_12(self):
# 1000 kt to psf
# truth value from NASA RP 1046
Value = A.cas2dp(1000, press_units='psf')
Truth = 5201.59
self.assertTrue(RE(Value, Truth) <= 2e-5)
def test_10(self):
# 1000 mph to psf
# truth value from NASA RP 1046
Value = A.cas2dp(1000, press_units='psf', speed_units='mph')
Truth = 3719.98
self.assertTrue(RE(Value, Truth) <= 2e-5)
def test_09(self):
# 1000 mph to in HG
# truth value from NASA RP 1046
Value = A.cas2dp(1000, press_units='in HG', speed_units='mph')
Truth = 52.5970
self.assertTrue(RE(Value, Truth) <= 2e-5)
def test_08(self):
# 244 kt to pa
# truth value from NASA RP 1046
Value = A.cas2dp(244, press_units='pa')
Truth = 9983.7
self.assertTrue(RE(Value, Truth) <= 2e-5)
def test_07(self):
# 280 kt to mm HG
# truth value from NASA RP 1046
Value = A.cas2dp(280, press_units='mm HG')
Truth = 99.671
self.failUnless(RE(Value, Truth) <= 2e-5)
def tas2ssec(tas,
alt,
oat,
ias,
speed_units=default_speed_units,
alt_units=default_alt_units,
temp_units=default_temp_units,
press_units=default_press_units):
"""
Return static source position error as speed error, pressure error and altitude error at sea level.
Returns delta_Vpc, delta_Ps, delta_Hpc and Vc
delta_Vpc = error in airspeed = calibrated airspeed - indicated airspeed corrected for instrument error
delta_Ps = error in the pressure sensed by the static system = pressure sensed in the static system - ambient pressure
delta_Hpc = altitude error at sea level = actual altitude - altitude sensed by the static system
Vc = calibrated airspeed at the test point
"""
P0 = U.press_conv(constants.P0, from_units='pa', to_units=press_units)
cas = A.tas2cas(tas,
alt,
oat,
speed_units=speed_units,
alt_units=alt_units,
temp_units=temp_units)
delta_Vpc = cas - ias
qcic = A.cas2dp(ias, speed_units=speed_units, press_units=press_units)
qc = A.cas2dp(cas, speed_units=speed_units, press_units=press_units)
delta_Ps = qc - qcic
delta_Hpc = SA.press2alt(P0 - delta_Ps, press_units=press_units)
# print 'Vc = %.1f, delta_vpc = %.1f, delta_ps = %.1f, delta_hpc = %.0f' % (cas, delta_Vpc, delta_Ps, delta_Hpc)
return delta_Vpc, delta_Ps, delta_Hpc, cas
def WOT_speed_orig(prop, altitude, weight = 1800, rpm = 2700, temp = 'std', \
temp_units = 'C', rv = '8', wing_area = 110, speed_units = 'kt' , \
MP_loss = 1.322, ram = 0.5, pwr_factor=1, mixture='pwr', \
wheel_pants=1):
"""
Returns the predicted speed at full throttle.
The MP_loss is the MP lost in the induction tract at full throttle at 2700
rpm at MSL. The default value is from the Lycoming power charts.
The ram is the percentage of available ram recovery pressure that is
achieved in the MP.
"""
press = SA.alt2press(altitude, press_units='in HG')
MP_loss = MP_loss * (rpm / 2700.)**2
cas_guess = 300
error = 1
if mixture == 'econ':
pwr_factor *= .86 # average power ratio to max power when at 50 deg LOP mixture
elif mixture == 'pwr':
pass
else:
raise ValueError('mixture must be one of "pwr" or "econ"')
while error > 0.0001:
dp = A.cas2dp(cas_guess, press_units='in HG')
# print 'dp =', dp
ram_press = ram * dp
# print 'Ram rise =', ram_press
MP = press - MP_loss + ram_press
pwr = IO.pwr(rpm, MP, altitude, temp=temp,
temp_units=temp_units) * pwr_factor
# print 'MP =', MP, 'Power =', pwr
tas = speed(prop, altitude, weight, pwr, rpm, temp = temp, \
temp_units = temp_units, rv = rv, wing_area = wing_area, \
speed_units = speed_units, wheel_pants = wheel_pants)
cas = A.tas2cas(tas, altitude, temp=temp, temp_units=temp_units)
error = M.fabs((cas - cas_guess) / cas)
cas_guess = cas
# print 'CAS =', cas, 'TAS =', tas
# print "MP = %.3f" % MP
# print "Pwr = %.2f" % pwr
return tas
def WOT_speed_orig(prop, altitude, weight = 1800, rpm = 2700, temp = 'std', \
temp_units = 'C', rv = '8', wing_area = 110, speed_units = 'kt' , \
MP_loss = 1.322, ram = 0.5, pwr_factor=1, mixture='pwr', \
wheel_pants=1):
"""
Returns the predicted speed at full throttle.
The MP_loss is the MP lost in the induction tract at full throttle at 2700
rpm at MSL. The default value is from the Lycoming power charts.
The ram is the percentage of available ram recovery pressure that is
achieved in the MP.
"""
press = SA.alt2press(altitude, press_units = 'in HG')
MP_loss = MP_loss * (rpm / 2700.) ** 2
cas_guess = 300
error = 1
if mixture == 'econ':
pwr_factor *= .86 # average power ratio to max power when at 50 deg LOP mixture
elif mixture == 'pwr':
pass
else:
raise ValueError, 'mixture must be one of "pwr" or "econ"'
while error > 0.0001:
dp = A.cas2dp(cas_guess, press_units = 'in HG')
# print 'dp =', dp
ram_press = ram * dp
# print 'Ram rise =', ram_press
MP = press - MP_loss + ram_press
pwr = IO.pwr(rpm, MP, altitude, temp = temp, temp_units = temp_units) * pwr_factor
# print 'MP =', MP, 'Power =', pwr
tas = speed(prop, altitude, weight, pwr, rpm, temp = temp, \
temp_units = temp_units, rv = rv, wing_area = wing_area, \
speed_units = speed_units, wheel_pants = wheel_pants)
cas = A.tas2cas(tas, altitude, temp = temp, temp_units = temp_units)
error = M.fabs((cas - cas_guess) / cas)
cas_guess = cas
# print 'CAS =', cas, 'TAS =', tas
# print "MP = %.3f" % MP
# print "Pwr = %.2f" % pwr
return tas
def tas2ssec2(tas,
ind_alt,
oat,
ias,
std_alt=0,
speed_units=default_speed_units,
alt_units=default_alt_units,
temp_units=default_temp_units,
press_units=default_press_units):
"""
Return static source position error as speed error, pressure error and
altitude error at sea level using speed course method.
Returns delta_Vpc, delta_Ps and delta_Hpc
tas = true airspeed determined by speed course method, or GPS
ind_alt = pressure altitude, corrected for instrument error
oat = outside air temperature, corrected for instrument error and ram temperature rise
ias = indicated airspeed, corrected for instrument error
std_alt = altitude to provide delta_Hpc for
delta_Vpc = error in airspeed = calibrated airspeed - indicated airspeed
corrected for instrument error
delta_Ps = error in the pressure sensed by the static system = pressure
sensed in the static system - ambient pressure
delta_Hpc = altitude error at std_alt = actual altitude - altitude
sensed by the static system
Uses analysis method from USAF Test Pilot School. Unlike some other
methods (e.g.. that in FAA AC 23-8B, or NTPS GPS_PEC.xls), this method
provides an exact conversion from TAS to CAS (some other methods assume
CAS = EAS), and it accounts for the effect of position error of altitude
on the conversion from TAS to CAS (some other methods assume pressure
altitude = indicated pressure altitude).
"""
tas = U.speed_conv(tas, speed_units, 'kt')
ind_alt = U.length_conv(ind_alt, alt_units, 'ft')
oat = U.temp_conv(oat, temp_units, 'C')
M = A.tas2mach(tas, oat, temp_units='C', speed_units='kt')
if M > 1:
raise ValueError('This method only works for Mach < 1')
delta_ic = SA.alt2press_ratio(ind_alt, alt_units='ft')
qcic_over_Psl = A.cas2dp(ias, speed_units='kt',
press_units=press_units) / U.press_conv(
constants.P0, 'pa', to_units=press_units)
qcic_over_Ps = qcic_over_Psl / delta_ic
Mic = A.dp_over_p2mach(qcic_over_Ps)
delta_mach_pc = M - Mic
if Mic > 1:
raise ValueError('This method only works for Mach < 1')
deltaPp_over_Ps = (1.4 * delta_mach_pc * (Mic + delta_mach_pc / 2)) / (
1 + 0.2 * (Mic + delta_mach_pc / 2)**2)
deltaPp_over_qcic = deltaPp_over_Ps / qcic_over_Ps
delta_Hpc = SA.alt2temp_ratio(
std_alt, alt_units='ft') * deltaPp_over_Ps / 3.61382e-5
# experimental - alternate way to calculate delta_Hpc that gives same answer
Ps = SA.alt2press(ind_alt, alt_units='ft', press_units=press_units)
delta_Ps = deltaPp_over_Ps * Ps
P_std = SA.alt2press(std_alt, alt_units='ft', press_units=press_units)
deltaPs_std = deltaPp_over_Ps * P_std
delta_Hpc2 = SA.press2alt(P_std - deltaPs_std,
press_units=press_units) - std_alt
delta_std_alt = SA.alt2press_ratio(std_alt, alt_units='ft')
asl = U.speed_conv(constants.A0, 'm/s', 'kt')
delta_Vpc_std_alt = deltaPp_over_Ps * delta_std_alt * asl**2 / (
1.4 * ias * (1 + 0.2 * (ias / asl)**2)**2.5)
actual_alt = SA.press2alt(Ps + delta_Ps,
press_units=press_units,
alt_units='ft')
cas = A.tas2cas(tas,
actual_alt,
oat,
speed_units='kt',
alt_units='ft',
temp_units='C')
return delta_Vpc, delta_Ps, delta_Hpc, cas
def tas2ssec2(tas, ind_alt, oat, ias, std_alt = 0, speed_units=default_speed_units, alt_units=default_alt_units, temp_units=default_temp_units, press_units = default_press_units):
"""
Return static source position error as speed error, pressure error and
altitude error at sea level using speed course method.
Returns delta_Vpc, delta_Ps and delta_Hpc
tas = true airspeed determined by speed course method, or GPS
ind_alt = pressure altitude, corrected for instrument error
oat = outside air temperature, corrected for instrument error and ram temperature rise
ias = indicated airspeed, corrected for instrument error
std_alt = altitude to provide delta_Hpc for
delta_Vpc = error in airspeed = calibrated airspeed - indicated airspeed
corrected for instrument error
delta_Ps = error in the pressure sensed by the static system = pressure
sensed in the static system - ambient pressure
delta_Hpc = altitude error at std_alt = actual altitude - altitude
sensed by the static system
Uses analysis method from USAF Test Pilot School. Unlike some other
methods (e.g.. that in FAA AC 23-8B, or NTPS GPS_PEC.xls), this method
provides an exact conversion from TAS to CAS (some other methods assume
CAS = EAS), and it accounts for the effect of position error of altitude
on the conversion from TAS to CAS (some other methods assume pressure
altitude = indicated pressure altitude).
"""
tas = U.speed_conv(tas, speed_units, 'kt')
ind_alt = U.length_conv(ind_alt, alt_units, 'ft')
oat = U.temp_conv(oat, temp_units, 'C')
M = A.tas2mach(tas, oat, temp_units='C', speed_units='kt')
if M > 1:
raise ValueError, 'This method only works for Mach < 1'
delta_ic = SA.alt2press_ratio(ind_alt, alt_units='ft')
qcic_over_Psl = A.cas2dp(ias, speed_units='kt', press_units=press_units) / U.press_conv(constants.P0, 'pa', to_units=press_units)
qcic_over_Ps = qcic_over_Psl / delta_ic
Mic = A.dp_over_p2mach(qcic_over_Ps)
delta_mach_pc = M - Mic
if Mic > 1:
raise ValueError, 'This method only works for Mach < 1'
deltaPp_over_Ps = (1.4 * delta_mach_pc * (Mic + delta_mach_pc / 2)) / (1 + 0.2 * (Mic + delta_mach_pc / 2 )**2)
deltaPp_over_qcic = deltaPp_over_Ps / qcic_over_Ps
delta_Hpc = SA.alt2temp_ratio(std_alt, alt_units='ft') * deltaPp_over_Ps / 3.61382e-5
# experimental - alternate way to calculate delta_Hpc that gives same answer
Ps = SA.alt2press(ind_alt, alt_units='ft', press_units=press_units)
delta_Ps = deltaPp_over_Ps * Ps
P_std = SA.alt2press(std_alt, alt_units='ft', press_units=press_units)
deltaPs_std = deltaPp_over_Ps * P_std
delta_Hpc2 = SA.press2alt(P_std - deltaPs_std, press_units = press_units) - std_alt
delta_std_alt = SA.alt2press_ratio(std_alt, alt_units='ft')
asl = U.speed_conv(constants.A0, 'm/s', 'kt')
delta_Vpc_std_alt = deltaPp_over_Ps * delta_std_alt * asl**2 / (1.4 * ias * (1 + 0.2 * (ias / asl)**2)**2.5)
actual_alt = SA.press2alt(Ps + delta_Ps, press_units = press_units, alt_units = 'ft')
cas = A.tas2cas(tas, actual_alt, oat, speed_units='kt', alt_units='ft', temp_units='C')
return delta_Vpc, delta_Ps, delta_Hpc, cas
