Real Time Thrust Estimation and Display for Ski-Jump ...

ISABE-2017-22569

Real Time Thrust Estimation and Display for Ski-Jump Takeoff of Fighter Aircraft A. Gogoi & B. Balaji [email protected] [email protected] Aeronautical Development Agency Bengaluru, India

ABSTRACT The take-off of a naval fighter aircraft from an aircraft carrier is extremely challenging. Accurate knowledge of thrust is essential for achieving successful takeoff from the aircraft carrier . In the present work, a model is developed and implemented for real-time display of thrust during the take-off phase of a naval fighter aircraft. This model uses the available engine sensors data and aircraft data to estimate the thrust. The model does not require any additional instrumentation to estimate the thrust. This paper describes the modelling details along with quantification of the modelling and measurement uncertainties. Results are shown for a few ski-jump takeoff tests carried out from a Shore Based Test Facility. The results indicated that the estimated thrust is within 1-3% of the expected thrust obtained from the Performance Cycle Deck of the engine manufacturer.

Keywords: Engine; Thrust; Nozzle area-pressure method; Ski-Jump

NOMENCLATURE A8

Nozzle Area

AE8

Effective nozzle area

FAP

Nozzle area-pressure method

FG

Gross Thrust, pounds

FN

Net Thrust, pounds

IFT

In-flight thrust

Mach

Mach Number

N1

Engine fan speed

PCD

Performance Cycle Deck

PLA

Power Lever Angle

Ps3

Compressor discharge pressure, psi

TDP

Turbine Discharge Pressure, psi

UFG

Uncertainty in Gross thrust

WFT

Fuel-mass flow method

W1

Engine Air flow rate, lb/sec

W1R

Engine Corrected Air Flow Rate, lb/sec



Angle of Attack, degree

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1. INTRODUCTION Take-off from aircraft carrier is extremely challenging for a fighter aircraft. The available takeoff distance in an aircraft carrier is extremely short as compared to a land-based runway. Accurate thrust estimation during the take-off phase is essential for safe takeoff from an aircraft carrier. The aircraft carriers can be broadly classified into two categories as follows [9] 1.Short Takeoff But Arrested Landing (STOBAR) 2. Catapult Takeoff But Arrested Landing (CATOBAR) In the STOBAR type of aircraft carriers, takeoff within a short distance is achieved with either

Thrust

Vectoring of the engine exhaust jet or by means of an upward bend (ski-jump) at the end of the deck. In the CATOBAR type of aircraft carriers, takeoff within the available distance is achieved with the assistance of steam catapult. The takeoff from the aircraft carrier with the ski-jump depends upon the thrust provided by the engine and the upward bend at the end of the deck. Thus, accurate prediction of thrust during the ski-jump takeoff phase is extremely critical for the naval aircraft. Accurate prediction of thrust becomes even more critical for singleengine aircraft as compared to twin-engine aircraft. Thus, in the present work, a model is developed for estimating thrust during ski-jump takeoff phase of a naval fighter aircraft. The model is used to calculate and display the real-time thrust during take-off of the naval fighter aircraft.

2. THRUST MODEL DETAILS The nozzle area-pressure method (FAP) and fuel-mass flow method (WFT) are the two popular method of inflight thrust (IFT) estimation. However, these models are primarily used for post-flight thrust calculation [2,3,4,5] and are not suitable for real-time thrust display due to its computational and measurement requirements. In the present work, the nozzle area-pressure method with suitable modification for faster computation and use of lesser number of sensors is used for estimation of real-time thrust. Although the fuel flow method is considered to be more accurate than nozzle area – pressure method, comparison of these two methods have been performed at high altitudes only [2,3,4]. The gross thrust for a choked nozzle can be expressed by the following equation.

𝐹𝐺 = 𝐴8𝐸 𝑃07

2 𝛾 𝛾+1

𝛾 +1 𝛾 −1

𝑃∞ 1− 𝑃07

𝛾 −1 𝛾

2𝛾 ……………….. 1 𝛾−1

where FG A8E P07 P



is the gross thrust is the effective nozzle area is the total pressure at nozzle inlet is the ambient pressure and is ratio of specific heat.

As seen in equation (1), the mass flow rate through the nozzle is calculated based on effective nozzle area (AE8) and nozzle pressure while jet velocity is calculated based on ratio of nozzle pressure and ambient pressure. The nozzle pressure is derived from Turbine Discharge Pressure (TDP) after accounting for the losses in the afterburner. The nozzle effective area and afterburner pressure losses are obtained from correlations from engine ground runs. These correlations enable faster real-time thrust computation required for display in the cockpit.

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ISABE-2017-22569 Net thrust is obtained by subtracting inlet momentum from gross thrust. Inlet momentum is calculated from freestream air velocity, inlet mass flow rate and inlet pressure recovery. Inlet mass flow rate is obtained from the corrected fan speed while inlet pressure recovery is obtained either from wind tunnel data or derived from the method described in [8]. The net thrust can be expressed as 𝐹𝑁 = 𝐹𝐺 − 𝑊1 𝑉∞ --------------------------------------------------------------------------------------(2) where V is free-stream air velocity and 𝑊1 = 𝑊1𝑅

𝛿 𝜃

,𝛿 =

𝑃1 101325

,𝜃 =

𝑇1 288 .15

, 𝑊1𝑅 = 𝑓 𝑁1𝑅 , 𝑃1 = 𝑓(𝑊1𝑅, 𝑀𝑎𝑐ℎ, ) -----------------(3)

3. UNCERTAINTY ESTIMATION In any dynamic system, the uncertainties can be classified into two types, namely modelling uncertainty and measurement uncertainty. In the present work, the modelling and the measurement uncertainties are estimated towards accurate thrust estimation. 3.1 Modelling Uncertainty Estimation It is assumed that the performance cycle deck (PCD) or the engine specification model of the engine manufacturer is an accurate representation of the engine. The nozzle area and afterburner loss correlations are adjusted to accurately match the thrust of the PCD for the takeoff conditions comprising of a limited range of altitude, Mach number and ambient temperature. The difference in thrust estimated by the algorithm as compared to the PCD for Sea Level Mach 0.0 and Mach 0.10 cases are shown in figure 1 and 2 respectively. It can be seen that the difference between the algorithm and the PCD is only about -0.42 to -0.95 % at Sea Level Mach 0.0 and -0.4% to -0.7% at Sea Level Mach 0.10 condition. The modelling uncertainty is less than one percent and hence can be considered acceptable. 0.00 -0.10

Percentage Deviation from PCD %

-0.20 -0.30 -0.40 -0.50 -0.60 -0.70 -0.80 -0.90 -1.00

Engine Specification Model (PCD) Thrust

Figure 1: Comparison of Engine Specification Thrust with Computed Thrust for Sea Level Mach 0.0 Case

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0.00

Percentage Deviation from PCD (%)

-0.10 -0.20 -0.30 -0.40 -0.50 -0.60 -0.70 -0.80 -0.90 -1.00

Engine Specification Model (PCD) Thrust

Figure 2: Comparison of Engine Specification Thrust with Computed Thrust for Sea Level Mach 0.10 Case

3.2 Measurement Uncertainty Estimation The in-flight thrust estimation method requires eight input parameters for thrust estimation. These parameters are the Power Lever Angle (PLA), fan speed (N1), fan inlet temperature (T1), Nozzle area (A8), pressure altitude, Mach number, Turbine Discharge Pressure (TDP) and Turbine Exit Temperature (T5C). The Power Lever Angle (PLA) is used to derive the co-relation between the Nozzle Inlet Pressure (P07) and Turbine Discharge Pressure (TDP). The pressure loss from the turbine exit section to nozzle entry is found to increase with increase in the PLA. At afterburner PLA settings, the additional afterburner pressure losses are also accounted to obtain the nozzle inlet pressure P07. The pressure altitude is used estimate the ambient pressure (P) used in evaluating the gross thrust in equation (1). The Turbine Exit Temperature (T5C), along with the engine bypass ratio, is used for estimating the value of Nozzle temperature which is then used for calculation of the specific heat ratio (). Since the engine is a low-bypass ratio turbofan engine, the nozzle temperature is lower than the turbine exit temperature (T5C) due to the mixing of the cooler bypass air. For afterburner operations, the nozzle Inlet Temperature which is used for calculating , is considered to be a function of the PLA. The nozzle area A8 is used for calculating the nozzle effective area AE8 which is used in estimation of the gross thrust shown in equation (1). Apart from the nozzle area A8, the nozzle effective area AE8 also depends upon the nozzle pressure ratio and the power lever angle (PLA). The fan speed (N1) and fan inlet temperature (T1) are used to calculate the corrected inlet mass flow rate which in turn is used for calculating the inlet momentum shown in equation (2). The flight Mach number is used for estimating the free-stream air velocity used in calculation of inlet momentum as shown in equation (2). The flight Mach number is also used in estimation of the inlet pressure recovery which is used to calculate the inlet mass flow rate. The measured parameters are required to be accurate for accurate thrust estimation. The nominal accuracy of each of the measured parameters along with its impact on thrust is shown in table 1. It can be seen from the table that Turbine Discharge Pressure (TDP) and Nozzle Area (A8) has the highest influence on thrust. A one percent

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ISABE-2017-22569 (1%) error in TDP leads to about 1.47% error in thrust while a one percent (1%) error in A8 leads to about 0.7% error in thrust. Table 1 Sensitivity of Measured Parameters

Parameter

Range

Accuracy

PLA

0-131 degrees

1 degree

0-110 %

±0.20%

-0.0226 , +0.0358

-54 to 140 deg C

± 1 degree

+0.0298, -0.0166

0-100%

±2%

+1.3996, -1.3807

-2000 to 75000 feet

± 75 feet

+0.1339, -0.1507

0-2.0

± 0.004

-0.1380, +0.1507

0-7 bar

± 1%

+1.4752, -1.4647

0-1000 deg C

± 11 deg C

Negligible

Fan Speed, N1 Fan Inlet Temperature, T1 Nozzle Area, A8 Pressure Altitude Mach Number Turbine Discharge Pressure, TDP T5C

Effect on Thrust in percentage 0 at MAXAB (MAXAB is from 129-131)

The uncertainty in net thrust due to measurement uncertainty can be expressed as [4]: 𝑈𝐹𝑁 =

𝐶𝑖,1 𝑈1

2

+ 𝐶𝑖,2 𝑈2

2

……………… 4

where Ci is the influence coefficient of th measured parameter U. The measurement uncertainty in Net Thrust (UFN) from equation (4) due to these eight parameters is about +2.04%, - 2.02% in thrust. In order to estimate the actual error of the TDP sensor especially during the take-off of the aircraft, calibration of the sensor is carried out. In the calibration process, known value of input pressure is applied to the sensor and the corresponding output voltage is recorded. The pressure value corresponding to the output voltage is compared with the input value and the difference is recorded. The calibration chart of the TDP sensor is shown in figure 3. It can be seen from the figure that accuracy of the sensor is very high around the range of 50 psi. Incidentally, the expected TDP value during the ski-jump take-off is also around 50 psi. Considering an error of -0.35,+0.0% in TDP reading during ski-jump takeoff, the uncertainty in thrust due to the TDP sensor reduces to -0.5%, 0%. The overall measurement uncertainty in thrust estimation during ski jump take-off reduces to +1.41,1.48% and becomes significantly lower than the nominal uncertainty shown in table 1. Hence, it can be said that the TDP sensor is suitable for thrust estimation during ski-jump take-off and the present method is likely to produce accurate results.

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Figure 3 Calibration Chart of Turbine Discharge Pressure (TDP) Sensor

4. INTAKE PRESSURE RECOVERY ESTIMATION The thrust provided by the Performance Cycle Deck (PCD) of the engine is used for evaluating the thrust computed by the algorithm. Intake pressure recovery is an input to the Performance Cycle Deck (PCD) of the engine and wind tunnel data of intake is taken as the baseline. The intake pressure recovery is modified with the method described in [8] to represent engine parameters more accurately.

5. RESULTS DURING SKI-JUMP TAKEOFF The ski-jump takeoff of the naval fighter aircraft was carried out on a Shore Based Test Facility. A picture of the Shore Based Test Facility is shown in figure 4. The test facility resemble a typical STOBAR category aircraft carrier with identical takeoff distance.

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Figure 4 : Shore Based Test Facility The aircraft is placed at the Restraining Gear Site( RGS) prior to takeoff and the throttle is set to maximum afterburner power setting. Once the expected thrust is displayed in the cockpit, the restraining gear is released for the takeoff. The plot of thrust during ski-jump take-off for two representative flights is shown in figure 5a while the plot of aircraft parameters like Altitude, Mach number and angle of attack () is shown in figure 5b. It can be seen from the figure 5b that there is a sudden jump in the angle of attack () as the aircraft leaves the ramp. It can be seen from the figure 5a that estimated thrust is very close to the expected thrust from the PCD. The initial discrepancy in thrust is due to the time taken by the afterburner to light-up as the throttle is moved to maximum afterburner for the ski-jump. The effect of angle of attack can be seen in the figure and a few percentage drop in thrust is observed at high angle of attack as the aircraft leaves the ramp. The increase in thrust with increase in forward speed can be seen in the figure. The deviation of estimated thrust computed by the algorithm from the PCD values is shown in figure 6 and the estimated thrust is within 1-3% of the expected PCD thrust for the two flights. The plot of turbine discharge pressure (TDP) for the two flights is shown in figure 7. It can be seen from the figure that the measured values are very close to the PCD values. Further, it can be seen that the TDP values during the takeoff phase is in the range of 45-50 psi and the measurement error at these readings is negligible as shown in figure 2. The comparison of Ps3, fuel flow rate and nozzle area during take-off are shown in figures 8, 9 and 10 respectively. It can be seen from the figure that these parameters agree well with the expected values obtained from the engine PCD. The engine internal parameters follow the same trend of the thrust computed by the algorithm.

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ISABE-2017-22569 The difference between the engine internal parameters like Ps3, TDP, A8 and fuel flow rates with respect to the engine model specification data for the two ski-jump takeoff is shown in figure 11. It can be seen from the figure that the deviation of the engine internal parameters are within 1-3% of the engine model specification.

Flight 1

Flight 2 Figure 5a Plot of Thrust during Ski-Jump Takeoff Alt/10

AOA

Mach*100

, Mach , Altitude

MACH*100

, Mach, Altitude

AOA

Time

Time

Flight 1

Flight 2

Figure 5b : Plot of Aircraft parameters during Ski-Jump Takeoff

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Altitude/10

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Flight 1

Flight 2 Figure 6 Plot of Thrust deviation during Ski-Jump Takeoff

Flight 1

Flight 2

Figure 7 Plot of Turbine Discharge Pressure (TDP) during ski-jump takeoff

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Flight 1

Flight 2

Figure 8 Plot of Compressor Discharge Pressure (ps3) during ski-jump takeoff

Flight 1

Flight 2 Figure 9 Plot of Fuel Flow Rate during ski-jump takeoff

Flight 1

Flight 2 Figure 10 Plot of A8 during ski-jump takeoff

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Percentage Deviation (%)

8 6 4

Ps3 Devistion TDP Deviation A8 Deviation Fuel Flow Deviation

2 0

-2

10 8 Percentage Deviation (%)

10

Ps3 Deviation TDP Deviation

6

A8 Deviation

4

Fuel Flow Rate Deviation

2 0

-2

-4

-4

-6

-6

-8

-8

-10

-10 Time

Time

Flight 1

Flight 2

Figure 11 : Deviation of measured parameters from Engine Model Specification

6. CONCLUSION A model is developed and implemented to estimate thrust during ski-jump take-off of a naval fighter aircraft. The modelling and measurement uncertainties are significantly reduced and the method is found to estimate thrust during take-off with good accuracy. The engine internal parameters like Ps3, TDP, A8, fuel flow rates agree well with the expected values of the engine specification model.

REFERENCES [1] Performance Cycle Deck of Engine Manufacturer [2] DONALD L HUGHES, “Comparison of Three Thrust Calculation Methods Using In-flight Thrust Data”, NASA TM 81360, July 1981. [3] TIMOTHY R CONNERS, “Measurement Effects on the Calculation of In-flight Thrust for an F404 Turbofan Engine” NASA TM-4140, September 1989. [4] RONALD J RAY, “Evaluation of Various Thrust Calibration Techniques on an F404 Engine” NASA TP 3001, April 1990. [5] RONALD J RAY, “Evaluating the Dynamic Response of In-flight Thrust Calculation Techniques During Throttle Transients” , NASA TM 4591, June 1994. [6] "Uncertainty in in-flight Thrust Determination, AIR-1678", SAE, August 2002. [7] THANH C VAN & RAYMOND S CISZEK, "Real Time Engine Thrust Calculation for Modern Fighter Aircraft", AIAA Paper 2004-4084. [8] ROHIT VASHISTHA & A. GOGOI, “Pressure Recovery Estimation of Single Engine Fighter Aircraft Air Intake”, ISABE Paper-2013-1404. [9] JOSE-LUIS HERNANDO & RODRIGO MARTINEZ-VAL, "Carrier Suitability of Land Based Aircraft", 28th International Congress of the Aeronautical Sciences, September 2012, Australia.

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