1 Title of the Invention
Pressure error correction for Air data parameters to improve accuracies in fighter aircraft
2 Field of the Invention
The present invention relates to avionics in general and in particular, to, method for improvement in accuracy of air data parameters obtained from pitot static probe in real time. The improvement in accuracies are achieved by estimation of pressure errors caused due to variations in flow field from free stream static pressureand correction of measured pressure data.
3 Background of the Invention
The air datasystem of an aircraft is a system in which total pressure created by forward motion of aircraft and static pressure of atmosphere surrounding it are sensed by pitot static probes and measured in terms of speed, altitude and rate of change of altitude (vertical speed). The system consists of a pitot static probe that is connected to air data system. The air data system obtains pressure data from pitot probes and converts pressure data to voltage. The air data parameters are used to obtaining the airspeed, altitude and vertical speed of aircraft that are used for primary navigation.
Sensing of total, or pitot pressure, and of static pressure is effected by probe, which is suitably located in airstream and measure these pressure. The accuracy of measured speed and altitude by means of pitot static probe depend on variations in air flow field from free stream static pressure due to probe itself and aircraft. These effects of such disturbances are greatest on static pressure measuring section of pitot static system giving rise to pressure errors which is defined as the amount by which local static pressure at a given point in the flow field.
The pressure errors are to be compensated using Pressure error corrections computed in real-time to improve accuracies of the air data

parameters. The pressure errors are specific to each type of aircraft and varies with atmospheric condition and highly dynamic. No general model can be used for computing pressure error correction.
The present invention describes a method for computation of pressure error correction in real time as a function of ratio of static pressure and total pressure that improves the accuracy of altitude and airspeed and mach number.
4 Brief Summary of the Invention
The definition of the aircraft state in terms of attitude and position at any instant of time is a highly critical aspect. The steady or quasi steady elements such as ambient pressure and velocity of the air relative to aircraft are obtained from Air data system (101). The parameters of aircraft position like speed in Indicated air speed, calibrated air speed, True air speed, Mach number, altitude and vertical velocity are obtained from Air data system (101) and referred as air data parameters. The aircraft performance and handling characteristics of aircraft are a function of the air data parameters. The air data parameters are to be measured and indicated with high accuracy as any error in these parameters can have flight safety concerns considering that aircraft flight envelope and handling are function of air data parameters sensed from ambient atmospheric condition in which the aircraft is flying.
The Air data system uses static port (201) and dynamic port (202) of pitot probes (102) (Fig. 2) to sense the static pressureand dynamic pressure and Total temperature probe (103) for impact temperature. The pressure sensors are vulnerable to errors considering their position on aircraft and the air flow disturbances around the pitotprobe(102) due to aerodynamic characteristics of aircraft. The errors in the pressured sensed by the pitot probe inlet and system error caused by transmission of the sensed pressure are referred as Pressure error. The pressure errors are to be corrected to eliminate the error caused by the probe due to aerodynamic

characteristics of aircraft and transmission error to enable indicate accurate values of air data parameters to the pilot and termed Pressure Error Correction. The pressure errors differ for each aircraft and thus Pressure error correction is to be modeled for each aircraft considering its unique aerodynamic behavior and system errors.
The present invention describes a method for generating model for pressure error correction of fighter aircraft and estimate air data parameters to improve the accuracies of the air data parameters to enable safe flight, accurate navigation and effectively perform desired mission.
5 Detail Description of the Drawings
Figure 1 shows the position and location of the pitot static probe (102), Total temperature probe (103 and Air data system (101) on the aircraft.
Figure 2: Error! Reference source not found.shows the schematic layout of Static port and Dynamic port of Pitot static probe that measures the static pressure and dynamic pressure.
Figure 3:shows Impact temperature measured as a function of measured voltage representing total temperature (UT)
Figure 4:shows uncorrected static pressure as a function of measured voltage representing Fine static pressure (Usf)
Figure 5:shows Pressure Error Correction derived from pressure data points measured from flight at different altitude and speed as a function of ratio of Static pressure to Total pressure
6 Detail Description of the Invention
The present invention particularly describes a method for improving accuracies of the air data parameters by eliminating the pressure errors due to the pitot probe and system. Conventionally, the pressure errors are modelled with respect to altitude and speed and fixed corrections are used. The specification describes a method for estimation of pressure

errors from air data system in real-time depending on flight conditions are used to correct air data parameters.
The Air Data Computations (ADC) in the present fighter aircraft is carried out by Mission computer in both primary mode and reversionary mode. True Air Speed (TAS), Indicated Air Speed (IAS), Mach number., Static Temperature, Standard Altitude, Baro Height, these are computed by taking the input data from the Air data system (101), total temperature probe and pilot control panel. The Pitot probe (102) is located in the aircraft forward equipment compartment and Temperature probe (103) on Port side of aircraft front fuselage (Fig. 1).
The Air data system(101) is an electro-mechanical device, producing DC voltages corresponding to the static pressure (PS), differential Pressure (QC) by taking input from pitotprobe (102) (Fig.2). The Air data system (101)measures and distributes static pressure (PS) and differential pressures (QC) in the form of DC voltages. It is also connected toTemperature probe (103) and the unit generates a DC voltage proportional to square root of Impact temperature (VTi) expressed in degrees kelvin. The Air data system outputs are used by Mission computer for air data computations.
To compute True Air Speed (TAS) data, in addition to differential and static pressures, an input equivalent to the outside air temperature is also required. This is obtained from Total temperature probe (103), which is located on the Port side of the front fuselage.
The Air data system (101) measures static and dynamic pressure, and total temperature. Input voltages representing these quantities are processed to compute pressure altitude, true and calibrated airspeeds, Mach number, static temperature, and air density ratio. The air data system provides the analog DC voltage outputs proportional to differential (QC) and static pressure (PS) pressures (0 to + 10V) and square root of impact temperature (-10 to +10V) respectively to the mission computer

which computes standard altitude, baro height, Mach No., TAS, IAS and static temperature. The Impact temperature, uncorrected static and dynamic pressure are computed from input sensor voltages. In case of static pressure, a combination of coarse and fine signals is used. Discretes associated with static and dynamic pressure inputs determine the validity of these quantities.
6.1 Impact temperature, Static pressure and Dynamic pressure
The impact temperature (Ti) computation from measured voltage representing total temperature (UT) is estimated by,
Tj = (0.310625 UT + 17.70625)2
The impact temperature measured as a function of the measured voltage is shown in Figure 3.
The Course static pressure(Psc) is estimated from measured voltage representing Course static pressure(Usc) by,
PSC=110USC
The air data parameters are safety critical parameters and hence validity of the data from source is ensured. Upon verification of validity of the data from Air data system (101), Uncorrected static pressure (Psi) is estimated. The uncorrected static pressure is estimated from the measured voltage representingFine static pressure (Usf)from Air data system (101)and varies as a function of the course static pressure. The characteristics of variation of Uncorrected static pressure (Psi) and course static pressure (Psc) is as follows,
If Psc> 41.25 Usf+ 412.5 then
Psj= 55 Usf+ 550
if Psc> 20.65 Usf + 206 then

Psi = 27.5 Usf + 275
if Psc> 12.75 Usf + 78.5 then
PSJ= 13.8 Usf+137
If course static pressure is not in the above defined band then, Psi = 11.7Usf+20
The uncorrected static pressure as a function of measured voltage representing Fine static pressure (Usf)is shown in Figure 4.
Similarly, uncorrected Dynamic pressure (Qc) is estimated from measured voltage representing Dynamic pressure (Uq) by
Qcj = 150 Uq
Thus, the impact temperature (Ti), Uncorrected static pressure (Psi) and uncorrected Dynamic pressure (Qc) are estimated from the voltages obtained from Air data system (101). Upon power on, the filtered difference between static temperature and Indian reference atmosphere value (5Ts) is initialised to zero.
6.2 P ress u re E rror C or recti o n
Pressure error Corrections are applied to both static and dynamic pressure to allow for static pressure error. These corrections are a function of the ratio of static to total pressure. Total pressure is the sum of the Uncorrected static and dynamic pressure
Pt=Psi + Qci
The pressure error in the pitot probe (102) varies with the speed, heading and roll of the aircraft as the measured pressures will depend on the flow of air over the probe that in turn is affected by the angle of incidence of the

air on the probe. The pressure error for each aircraft varies with respect to its pitot position and manoeuvring of aircraft.
The pressure error for the described air data system (101) installed on the specific aircraft is obtained by flight test methodology. The aircraft is flown at various speeds and altitude as per its flight envelope along with another type of aircraft that is calibrated for its air data parameters and validated. The altitude and speed values of both aircraft are recorded through instrumentation and within Mission computer and the data analysed to derive a correction factor based on the difference between the air data parameters from the reference aircraft.
The pressure error correction must cover the full range of conditions and variables for which it will be applied and is achieved by making measurement of ambient pressure at the aircraft independently of aircraft system. The pressure error is derived by comparing difference between the static pressure recorded on own aircraft and pacer aircraft whose air data parameters are calibrated and validated. The input pressure and temperature sensed by the probes that are converted to voltage by the Air data system is recorded on both aircraft. All relevant variables that are used for computation of airspeed and altitude are also recorded on both aircraft for data analysis and comparison. The own aircraft and the pacer aircraft are equipped with on-board Inertial navigation system that prove speed, altitude and attitude of the aircraft obtained from inertial system and Global positioning system.
The two aircraft operate in formation with a separation of at least 40 feet and in steady level flight and same speed throughout the test flight envelope. The method involves correlation of the air data parameters at the same given particular instant of time and is time synchronised. The flight data comprising the static pressure, dynamic pressure and total temperature that are raw signals converted to voltage by air data sytem and the derived parameters such as impact temperature, uncorrected

static pressure, and uncorrected dynamic pressure are recorded in both variants of aircraft along with time stamp. The air data parameters are compared with the speed and altitude from the Inertial Global position system and the difference between the parameters for both aircraft is computed. Considering high accuracy of the hybrid altitude obtained from Inertial Global position system, both aircraft fly with the speed, altitude and . attitude with reference to Inertial Global position system. The pacer aircraft provides ambient pressure as a reference for that sensed by the subject aircraft. Comparison of pitot pressures gives a means of estimating the pressure errors and correction of pitot system of the subject aircraft. The difference in the pressure values between the two aircraft are computed and a equation representing the correction is derived.
The pressure error corrections (c) derivedbased on flight test results and comparison with pacer aircraft is as follows,
When(Psj / Pt) < 0.33 thenc = 0.032 - 0.0312 (Psi/ Pt)
When(Psj / Pt) < 0.424 thenc = 0.052 - 0.094 (Psi/ Pt)
when(Psj / Pt) < 0.515 thenc = 0.1052 - 0.219 (Psj/ Pt)
when(Psj / Pt) < 0.54 thenc = 0.163 - 0.26 (Psi/ Pt)
when(Psj / Pt) < 0.896 thenc = - 0.004. + 0.008 (PSj/ Pt)
The Pressure Error Correction derived from pressure data points measured from flight at different altitude and speed as a function of ratio of Static pressure to Total pressure is shown in Figure 5..
The dynamic pressure is computed as difference between Total pressure and Static pressure, Qc = Pf- Ps

6.3 Pressure altitude, Calibrated airspeed, and Mach number
Pressure altitude, calibrated airspeed, and Mach number are computed from static and dynamic pressure in accordance with the ISA relationships. In the case of calibrated airspeed and Mach number, a quadratic approximation is used for supersonic speeds. This approximation was derived as a best fit to the exact relationship over the ranges Vc/ao = 1 to 1.8 and M = 1 to 1.8 respectively. The Pressure altitude (Hp)varies with reference to static pressure and varies as per the lapse rates defined in International Standard Atmosphere and the relationship is as follows,
If (Ps> 226.321) then Hp = 145442.16Tl -(Ps /P0)° ,902631
if (Ps > 54.749) thenl-lp = 36089.24 - 20805.80 In (Ps / 226.321)
else, Hp= 65616.80- 710.79 Tl-(54.749/Ps)002927
The Calibrated air speed(Vc) is estimated as a function of the ratio of dynamic pressure and ISA static pressure at sea level (Po).
If Qc/P0< 0.89293 then
. Else
Vc = a0 (0.229 + 0.8118 VQC/P0 + 0.0043 Qc/P0)
The Mach number (M) is estimated as a function of the ratio of dynamic pressure and static pressure at flight level (Ps).
If w< 0.89293 then
M = J{s[{\+Qc/Psyn-\\Else

M = 0.229 + 0.8118 VQC / Ps + 0.0043 Qc / Ps
True airspeed (V) is computed as Mach number multiplied by the speed of sound (proportional to the square root of static temperature).
V = a0 M V(TS / T0) where, a0is ISA speed of sound at sea level

We Claim,
1. A method for computation of Pressure Error Correction in real time based on present ambient atmospheric and flight condition as a function of ratio of static pressure and total pressure to improve accuracy of altitude and airspeeds characterized in that:
- Air data parameters of air speedand altitude are obtained from Air data system (101), Static pressure probe (102) and Total temperature probe (103) (Figure 1); and
- Generating model for pressure error correction of subject fighter aircraft based on flight test with another pacer aircraft and estimate accurate air data parameters.
2. A method for computation of pressure error correction in real time as said in claim 1 wherein; the Air data system provides analog DC voltage outputs proportional to differential (QC) and static pressure (PS) pressures (0 to + 10V) and square root of impact temperature (-10 to +10V) respectively to Mission computer which computes standard altitude, baro height, Mach No., TAS, IAS and static temperature. The Impact temperature, uncorrected static and dynamic pressure are computed from input sensor voltages
3. A method for computation of pressure error correction in real time as said in claim 1 and 2 wherein;
- combination of coarse and fine signals is used for estimation of Static pressure;
- impact temperature (Ti) computation from measured voltage representing total temperature (UT);
- uncorrected static pressure is estimated from the measured voltage representing Fine static pressure (Usf) from Air data system (101) and varies as a function of course static pressure;and

- uncorrected Dynamic pressure (Qc) is estimated from measured voltage representing Dynamic pressure (Uq).
4. A method for computation of pressure error correction is modeled based on flight test as said in claim 1 wherein;
- The subject aircraft is flown at various speeds and altitude as per its flight envelope along with another pacer aircraft that is calibrated for its air data parameters and validated;
- The altitude and speed values of both aircraft are recorded through instrumentation and within Mission computer and the data analysed to derive a correction factor based on difference between air data parameters from reference pacer aircraft;
- The two aircraft operate in formation with a separation of at least 40 feet and in steady level flight and same speed throughout the test flight envelope using navigation parameters from onboard Inertial Global position system;
- The air data parameters are compared with the speed and altitude from the Inertial Global position system and the difference between .the parameters for both aircraft is computed along with time stamp; and
- The Pressure Error Correction derived from pressure data points measured from flight at different altitude and speed as a function of ratio of Static pressure to Total pressure.