Engine Power Settings Research Articles

On Hysteresis In A Variable Pitch Fan Transitioning To Reverse Thrust Mode And Back

Abstract A novel hysteresis phenomenon during the transition to and back from the reverse thrust mode in a Variable Pitch Fan (VPF) is identified and characterised in this work. This is done by using a three-dimensional (3D) fully transient Unsteady Reynolds averaged Navier?Stokes (URANS) with the transitioning fan blade aerofoils simulated by an adaptation of the mesh displacement method. The VPF is modelled to be transitioning in a modern 40000 lbf geared high bypass ratio turbofan engine architecture at "Approach Idle" engine power setting in a typical twin-engine airframe with the flaps, slats, and spoilers set for an aircraft touchdown airspeed of 140 knots. The transition to reverse thrust mode involves flow starvation into the engine, formation of recirculation zones in the bypass duct and the establishment of the reverse stream, all of which occurs in the opposing presence of the free stream flow at aircraft touchdown velocity. The transition back to forward flow mode involves the gradual re-establishment of the free stream which is opposed by the presence of the reverse stream within the engine. It is quantified that in the transition to reverse thrust, the blockage develops with a larger time delay than the disappearance of the blockage during the transition back due to the interplay of the temporal dynamics of fan blade motion and flow field response. The hysteresis phenomena described in this work are critical in properly developing control schedules to adapt for potential bi-stable flow field development during the landing run.

On the Flow Physics During the Transition of a Variable Pitch Fan From Nominal Operation to Reverse Thrust Mode

Abstract The flow field during the transition of variable pitch fans (VPFs) from nominal operation to reverse thrust mode at typical “Approach Idle” engine power setting and aircraft touchdown speed of 140 knots is described in this work. An integrated airframe-engine-VPF research model that features a future 40,000 lbf geared high bypass ratio engine installed on a twin-engine airframe in landing configuration is used to explore the flow field in a fully transient unsteady Reynolds-averaged Navier–Stokes (URANS) simulation with imposed wall motion. A novel methodology that implements an adaptation of a mesh displacement equation to mimic the fan blade airfoil rotation is developed. The implementation of this method with gradual, small-step deformation along with an automated mesh update routine enables a high quality, near “real-time” simulation of the complete transition. The flow field during transition is characterized by the evolution from full typical forward flow through the engine to the development of massive recirculation regions at the feather pitch setting and finally to development of a reverse flow from the bypass nozzle to the fan passages. In the paper, the transient development of the various flow features through different stations of the engine flow path apropos the fan blade airfoil rotation to reverse thrust mode are discussed in detail. Also, the temporal development of the mass flow ingested through the engine, airframe decelerating force, and distorted flow at the core engine inlet are described. A hitherto unresolved fan power peaking during the middle of the transition and higher power requirement at reverse thrust mode is captured. The effect of fan rotational speed in terms of engine power setting and the aircraft touchdown velocity on the transition flow physics is explored. A comparison of this fully transient approach with discrete steady-state runs for different stagger angle settings is presented. The flow physics during transition to reverse thrust mode as described in this study is critical in understanding the feasibility of using VPF for reverse thrust in future aircraft. The new capability to study the transition in a fully transient simulation can be used as a design development aid to define design and control characteristics of the reverse thrust VPF.

Acoustic measurement for resin-based acoustic liner on turbofan engine testbed

Acoustic liner is one of the most effective devices for suppressing fan noise emitted by aircraft engines. IHI and Japan Aerospace Exploration Agency (JAXA) have developed a resin-based acoustic liner for fan noise attenuation that is lighter than conventional acoustic liners in production. In this study, an acoustic measurement of the full-scale resin-based acoustic liner was conducted on the turbofan engine testbed installed at the JAXA test facility to evaluate its noise reduction capability in actual operating environment. The in-duct and in-cell acoustic measurements were performed during the test campaign. A baseline hardwall liner and the production acoustic liner were also tested for comparison. The tonal and broadband fan noise reduction of the resin-based acoustic liner relative to the hardwall liner was evaluated as well as that of the production acoustic liner. According to the test results obtained in the enclosed testbed, it was found that the resin-based acoustic liner has good noise reduction performance as expected for typical acoustic liners. The maximum noise reduction is equivalent to the production acoustic liner for both tonal and broadband components over the wide range of engine power settings.

Powered Low-Speed Experimental Aerodynamic Investigation of an Over-Wing-Mounted Nacelle Configuration

Over-wing integration of ultrahigh bypass turbofan engines can be a solution for next-generation commercial transport aircraft since it eliminates the ground clearance issue and it has the potential to reduce ground noise due to acoustic shielding. Moreover, a unique characteristic of this installation type is the powered lift benefit at low-speed flight conditions. This paper aims to experimentally investigate the effect of the engine power setting on the low-speed aerodynamic performance of an over-wing-mounted nacelle configuration comprising a conventional tube-and-wing layout. Thus, low-speed wind tunnel tests were performed for a half-span powered scale model of the aforementioned configuration. The effect of the engine power setting on the wing lift and spanwise pressure distributions was investigated. The experiments were carried out for angles of attack varying from 0 to 6° and inlet mass flow ratios up to 2.4. The results were used to validate computational fluid dynamics simulations conducted for the same wind tunnel conditions. It has been shown that a significant powered lift benefit can be achieved for the studied configuration, without a penalty in the net propulsive force and that the lift increases linearly with the inlet mass flow ratio. Furthermore, it was observed that the engine power setting largely influences the pressure distributions along the wing, especially at the spanwise sections closer to the nacelle. The low-momentum zone created upstream of the engine at high power settings reduces the pressure at the wing’s upper surface, which is the main factor responsible for the increased lift. By taking advantage of such behavior, drag can potentially be reduced at takeoff and climb due to a lower flap setting required for the same lift.

Sensitivity Analysis of the Overwing Nacelle Design Space

The overwing nacelle (OWN) concept refers to aircraft designs where the engine is installed above the wing. The OWN configuration offers several advantages over conventional underwing nacelle (UWN) vehicles, which include improved fuel-burn and propulsive efficiencies due to the feasibility of ultrahigh-bypass-ratio turbofans, as well as reduced noise. However, a nonoptimal OWN design can result in large transonic drag penalties that can potentially outweigh the aforementioned benefits. We study the OWN design problem from an aerodynamics and propulsion perspective using the NASA Common Research Model, a notional 90,000-lb-thrust-class turbofan model, and Reynolds–averaged Navier–Stokes simulations. We first quantify the sensitivity of the drag, lift, and pressure recovery to variations in engine location and power setting; and we identify trends. Then, we perform aerodynamic design optimization of the wing and nacelle to determine OWN performance improvement from outer-mold-line refinement at a favorable engine installation location. A 20% reduction in drag is achieved for the optimized OWN configuration, highlighting the sensitivity of OWN aerodynamics to airframe contours. However, compared to the UWN baseline, the optimized OWN drag is 5% higher at the same lift and worsens significantly at higher lift.

Global warming potential optimization of a turbofan powered unmanned aerial vehicle during surveillance mission

Gas emissions from all sources including air transport contribute to global warming. There are tremendous efforts to decrease the greenhouse gas emissions in all sectors. In this study, for a turbofan engine powered unmanned aerial vehicle, Global Warming Potential (GWP) calculations have been performed covering a range of flight Mach number and altitude. Unmanned aerial vehicles have different mission profiles and usage characteristics than manned aircraft. Therefore their emissions are also different than manned aircraft due to different flight altitude, loitering (surveillance) speed and engine power setting etc. Global warming potential was calculated with a genuine code combining aircraft performance, engine performance and exhaust gas emissions for all flight points during surveillance mission. Exhaust gases of CO2, H2O and NOx were calculated and at each leg of flight as they contribute to GWP. Results were presented to illustrate the GWP magnitudes for different flight speed and altitude in a comparative manner for the turbofan powered UAV. Optimization calculations were done by using genetic algorithm. For an investigated range of 6–16 km flight altitude and 0.25–0.75 flight Mach number, total surveillance mission GWP is lowest around 13 km and 0.37 Mach number as optimum values. However, GWP per hour during constant speed and altitude surveillance (loiter) is minimum around 15 km and 0.43 Mach number for assumed UAV flight mission profile.

Predicting Lean Blowout and Emissions of Aircraft Engine Combustion Chamber Based on CRN

Abstract To predict the pollutant emissions and lean blowout, chemical reactor network (CRN) model is applied to the modern aircraft engine combustion chamber. In this study, the CRN which represent the major features of aerodynamics and combustion in the combustion chamber is set up based on the OpenFOAM simulation results. The boundary and the initial conditions used for the CRN derive from the operating modes of typical aircraft engine cycle. A 21 species 30 steps chemical mechanism of kerosene is employed in the CRN method. The levels of pollutant emissions are obtained under four ICAO engine power settings of idle, approach climb and take off. The lean blowout equivalent ratio is evaluated at the idle power setting. The results will be helpful to predict the aircraft engine exhaust emissions and lean blowout (LBO).

Nonintrusive Gas-Turbine Engine-Exhaust Characterization Using Acoustic Measurements

This paper reports exhaust measurements performed using a nonintrusive acoustic technique on a JT15D-1A turbofan research engine. Using multiple streamwise displaced microphones, the integrated exhaust velocity and static temperature were measured simultaneously in engine flows with bypass Mach numbers up to 0.48. A Kiel pressure and K-type thermocouple probe was used to assess the accuracy of the acoustic measurements. The Integrated velocity and static temperature rms errors of and 2.3 K were identified. Using probe-based measurements at a low engine power setting, two calibration constants were identified for conversion from integrated flow parameters to mass flow and thrust. Applying the calibration constants to the acoustic measurements, and 200 N rms errors in mass flow and thrust were observed. These promising results indicate that the acoustic technique may be used to characterize and monitor gas-turbine engine performance. To the authors’ knowledge, this is the first time a nonintrusive acoustic technique has been used to characterize engine flows with Mach numbers greater than 0.3.

Power optimization of a single propeller airplane take-off run on the basis of lateral maneuver limitations

The problem of selecting optimal take-off run power of a single propeller airplane on the basis of the lateral maneuver limitations is analyzed in this paper. The single propeller reactive torque as the function of engine power and speed generates non-aerodynamic asymmetric load disturbance. The capability of lateral flight control surfaces to compensate for this asymmetric load and perform the required maneuver defines the safe flight envelope boundaries in the form of engine power setting – airspeed relation. The take-off run length consisting of acceleration, decision making, brake activation and airplane stopping length distances is calculated for engine power settings along the safe flight boundaries defined by the lateral flight control capabilities. The minimum of take-off run along thus defined safe flight boundary determines the engine optimal power setting. The results are presented for the example of the single propeller airplane on concrete runway surface.

Longitudinal stability characteristics of the LASTA airplane

PurposeThis paper presents the longitudinal stability analysis of a single tractor propeller airplane LASTA at high engine power settings. This analysis is part of the ongoing process of certifying the airplane for civil use according to the civil regulations CS-23.Design/methodology/approachThe design methodology that is presented in the paper consists of comparing flight test aerodynamic and calculation results. The methods used here are standard and routinely used in flight testing.FindingsFlight testing results indicate that at low airspeeds the cumulative destabilizing effects because of high values of the angle of attack and high power settings are about 6 per cent of MAC. This value is in a very good agreement with published data.Practical implicationsThe information presented in this paper are new, and are very specific to this one aircraft configuration. The methods used here are standard and widely used in flight testing.Originality/valueThe information in this paper presents flight test results. There are not many publications in this area.

Framework for Predicting Noise–Power–Distance Curves for Novel Aircraft Designs

Along with flight profiles, noise–power–distance curves are the key input variable for computing noise exposure contour maps around airports. With the development of novel aircraft designs (incorporating noise-reduction technologies) and new noise-abatement procedures, noise–power–distance datasets will be required for assessing their potential benefit in terms of noise reduction around airports. Noise–power–distance curves are derived from aircraft flyover noise measurements taken for a range of aircraft configurations and engine power settings. Clearly then, empirical noise–power–distance curves will be unavailable for novel aircraft designs and novel operations. This paper presents a generic framework for computationally generating noise–power–distance curves for novel aircraft and situations. The new framework derives computationally the noise–power–distance noise levels that are normally derived experimentally, by estimating noise level variations arising from technological and operational changes with respect to a baseline scenario, where the noise levels are known or otherwise estimated. The framework is independent of specific prediction methods and can use any potential new model for existing or new noise sources. The paper demonstrates the methodology of the framework, discusses its benefits, and illustrates its applicability by deriving noise–power–distance curves for an unconventional approach operation and for a future concept blended wing–body aircraft.

Aerodynamic Effects of Propulsion Integration for High Bypass Ratio Engines

This work describes the assessment of the effect of engine installation parameters such as engine position, size, and power setting on the performance of a typical 300-seater aircraft at cruise condition. Two engines with very high bypass ratio and with different fan diameters and specific thrusts are initially simulated in isolation to determine the thrust and drag forces for an isolated configuration. The two engines are then assessed in an engine–airframe configuration to determine the sensitivity of the overall installation penalty to the vertical and axial engine location. The breakdown of the interference force is investigated to determine the aerodynamic origins of beneficial or penalizing forces. To complete the cruise study, a range of engine power settings is considered to determine the installation penalty at different phases of cruise. This work concludes with the preliminary assessment of cruise fuel burn for two engines. For the baseline engine, across the range of installed positions, the resultant thrust requirement varies by 1.7% of standard net thrust. The larger engine is less sensitive with a variation of 1.3%. For an assessment over a 10,000 km cruise flight, the overall effect of the lower specific thrust engine shows that the cycle benefits of in specific fuel consumption are supplemented by a relatively beneficial aerodynamic installation effect but offset by the additional weight to give a fuel-burn reduction.

Development of hearing protection device calculators for the F-35A and F-35B tactical aircraft

Simple spreadsheet-based near field noise prediction tools for the F-35A (conventional takeoff and landing) and F-35B (short takeoff and vertical landing) aircraft have been developed recently. Near field noise predictions are based on empirical data collected during static engine runs conducted at Edwards AFB, and during hover/vertical landing operations at MCAS Yuma in September, 2013. The tools are programmed to calculate near field one-third octave band sound pressure levels as a function of engine power setting and listener location. The calculated near field levels can be used to determine the Total Daily Exposure (TDE) of personnel required to work in the vicinity of powered aircraft. Alternatively, the calculators can be used to determine the amount of hearing protection required so that the personnel TDEs remain below 1.0. The calculators are adjustable so that they can be employed for differing domestic and international noise exposure regulations. [Work supported by the Air Force Research Laboratory and Ball Aerospace.]

Stability characteristics of the single-engine tractor propeller airplane in climb

This paper presents the analysis of effects on the airplane longitudinal stability reserve by the single tractor propeller at high engine power settings. The analysis was done within the terminal flight phases. The airplane has the original and modified horizontal tail shape. The horizontal tail modifications were applied to the second prototype of the “LASTA” airplane during the development of flight testing. The airplane stability problem was initiated by the rise of pilot average mass diverging in practice from the specification of 86 kg (190 lbs), as defined in aviation regulations and installation of measurement equipment. The “LASTA” airplane model is here analyzed for two different horizontal tail shapes by the combination of the calculating aerodynamics and flight testing methods. The paper makes the presentation of the introduction phase of the airplane design and construction process.

A gaseous emissions analysis of commercial aircraft engines during test-cell run

This study presents an analysis of the seventeen complete measurements of gaseous emissions of eleven different CFM56-7B26 turbofan engines during the regular test-cell runs. A novel measurement setup is established in which the sample probe is located immediately behind the low pressure turbine rotor. Recording a total of 49,190 data points for 26 raw variables (obtained from both the emission measurement and engine test systems), emissions are characterized with stable engine power settings representing main flight phases, as well as fuel flow of the engine during the entire run. Carbon monoxide and total hydrocarbons concentrations are found to be a strong function of fuel flow at low power settings, and nitrogen oxides emissions at high power settings. In comparison to emission indices provided by the International Civil Aviation Organization, the current study results in considerably higher carbon monoxide and hydrocarbons under idle conditions, with relatively lower nitrogen oxides at almost all of the power settings. On the other hand, significant differences in levels of nitrogen oxides are observed from one engine test to another, emphasizing the effects of the engine and the ambient air conditions. The average NO2/NOx are found to be 40% (±16) and 32% (±9) for ground and approach idles, respectively, while at high power settings, above N1 = 64%, the NO2/NOx ratios stabilize at around 5–6% (±1–2). Since the emissions were monitored and recorded during the entire engine run, the start and end moments (cold and hot idles) of the tests are compared. The average carbon monoxide and hydrocarbons concentrations are found to be 14% and 35% lower at hot ground idles compared to cold ground idle, with the results for nitrogen oxides and carbon dioxide being 22% and 0.5% higher. The effect of ambient air temperature on emissions are also investigated and quantified for certain temperature ranges. It has been determined that the hydrocarbons and carbon monoxide concentrations decrease with ambient air temperature, yet the nitrogen oxides and carbon dioxide concentrations increase. Findings related to ambient air temperature are found to corroborate the findings related to running at cold and hot idles, linking to a positive correlation between the ambient air temperature, the combustor inlet temperature and the exhaust gas temperature.

Technique for Reducing the Effects of Nonlinear Terms on Electric Field Measurements of Electric Field Sensor Arrays on Aircraft Platforms

AbstractA generalized technique has been developed that reduces the contributions of nonlinear effects that occur during measurements of natural electric fields around thunderstorms by an array of field mills on an aircraft. The nonlinear effects can be due to nearby charge emitted by the aircraft as it acquires and sheds charge, but the nonlinear effects are not limited to such sources. The generalized technique uses the multiple independent measurements of the external electric field obtained during flight to determine and remove nonlinear contaminations in the external vector electric field. To demonstrate the technique, a simulated case with nonlinear contaminations was created and then corrected for the nonlinear components. In addition, data from two different field programs utilizing two different aircraft and field mill configurations, each containing observable and different nonlinear effects, were also corrected for the significant nonlinear effects found in the field mill outputs. The expanded independent measurements in this new technique allow for the determination and correction of components in the field mill outputs from almost any measurable source. Alternate utilization of the technique can include removing effects in the aircraft charging such as aircraft altitude, cloud properties, engine power settings, or aircraft flap deployment. This technique provides a way to make more precise measurements of the true external electric field for scientific studies of cloud electrification.

Adaptation of the Beveled Nozzle for High-Speed Jet Noise Reduction

The noise of high-speed jets, generated by nozzle exhaust systems that resemble typical jet engine installations on fighter aircraft, has been investigated with the main objective of developing noise reduction designs. To this end, the aeroacoustic characteristics of beveled nozzles have been examined in model scale tests and their performance relative to a conventional round nozzle system has been established. Two different physical models with appropriate area ratios to simulate the engine operating conditions at different power settings were tested; these correspond to 96% N1 (Military or MIL power) and 91% N1 (cutback power). Aeroacoustic measurements and analyses have been carried out for nozzles at bevel angles of 20o, 24o, 28o and 35o for MIL power, and 24o and 32o for cutback power. For the same plenum conditions, the mass flow rates for the beveled nozzles are identical to those of the baseline round nozzle, for a wide range of NPR. The thrust coefficients are higher for the beveled nozzles for both the nozzle geometries at typical MIL power and cutback power, respectively. The increase in the thrust coefficient ranges from ~0.75% for the bevel35 to ~2% for bevel28, in spite of the slight deflection of the jet plume. The reasons for this phenomenon are examined. The beveled nozzles produce at least the same or greater absolute thrust as the baseline nozzle. The beveled nozzles specifically reduce the noise in the peak polar radiation angles; the maximum noise benefit is observed in the azimuthal direction of the longer lip. The magnitude of noise reduction increases with increasing bevel angle, with the largest reduction for bevel35. An examination of the dBA metric, which is used in noise exposure studies, indicates that there is a noise benefit of ~3 dBA to ~4 dBA in the peak radiation sector. The noise benefit in the azimuthal direction of the shorter lip is only slightly lower than that in the direction of the longer lip. I. Introduction he noise of high-speed jets, as from the engines that power fighter aircraft, is a major concern for both military personnel and civilians living close to military bases. Military jets are powered by turbojets or very low bypassratio turbofan engines (BPR ≤ 0.3). The attendant high jet velocities produce extremely high levels of jet noise. This problem is severe for operations on aircraft carrier decks because military personnel are stationed very close to the aircraft during the launch and the landing of carrier-based fighter aircraft. Pilots perform training missions, called Field Carrier Landing Practice, at military bases that mimic the actual flights at the appropriate engine power settings, thereby creating a huge noise problem for the surrounding communities. The magnitude of the problem is first highlighted with a comparison of the noise footprints from the F/A-18 A/B and the Boeing 737-800 aircraft. The noise foot print is calculated as follows: first, A-weighted spectra are calculated from the measured spectra over a wide range of radiation angles. The A-weighting adjusts the full-scale spectrum measured by the microphone to account for the response of the human ear, with the low frequency levels decreased by several decibels, and the high-frequency portion of the spectrum relatively unmodified. The energy levels in the resulting spectrum at the different frequencies are logarithmically summed to produce one dBA number at each angle. The maximum dBA

Minimizing Sampling Loss in Trace Gas Emission Measurements for Aircraft Engines by Using a Chemical Quick-Quench Probe

This paper describes the development and testing of a gas sampling probe that quenches chemical reactions by using supersonic expansion and helium dilution. Gas sampling probes are required for accurate measurement of exhaust emissions species, which is critical to determine the performance of an aircraft engine. The probe was designed through rounds of computational modeling and laboratory testing and was subsequently manufactured and then tested at the University of Tennessee Space Institute behind a General Electric J85 turbojet engine at different power settings: idle, maximum military, and afterburning. The experimental test results demonstrated that the chemical quick-quench (CQQ) probe suppressed the oxidation of carbon monoxide (CO) inside the probe system and preserved more CO at afterburning conditions. In addition, the CQQ probe prevented hydrocarbons from being partially oxidized to form CO at idle powers and measured higher hydrocarbons and lower CO emission compared with a conventional probe at that low power condition. The CQQ probe also suppressed nitrogen dioxide (NO2) to nitric oxide (NO) conversion through all engine power settings. These data strongly support the conclusion that the CQQ probe is able to quench unwanted chemical reactions inside the probe for all engine power levels.

Application of near-field acoustical holography to high-performance jet aircraft noise.

Near-field acoustical holography (NAH) measurements were made to visualize the sound field in the jet exhaust region of an F-22 Raptor. This is one of the largest-scale applications of NAH since its development in the 1980s. A scan-based holographic measurement was made using a 90-microphone array with 15-cm regular grid spacing for four engine power settings. The array was scanned through 93 measurement positions along three different measurement planes in a region near 7 m from the jet centerline and 23 m downstream. In addition, 50 fixed reference microphones were placed along the ground 11.6 m from the jet centerline, spanning 30.8 m. The reference microphones have been used to perform virtual coherence on the measurement planes. Statistically optimized NAH has been used to backpropagate the sound field to the source region for low frequencies and to identify jet noise characteristics. Ground reflection interference and other non-ideal measurement conditions must be dealt with. Details relating to jet coherence-lengths and their relation to reference microphone requirements will be discussed. Preliminary results of this ongoing work will be presented. [Work supported by Air Force SBIR.]

Emissions of Volatile Particulate Components from Turboshaft Engines Operated with JP-8 and Fischer-Tropsch Fuels

Particulate emissions from two types of helicopter turboshaft engines operated with military JP-8 and paraffinic Fischer-Tropsch (FT) fuels were characterized as an objective of the field campaign held at the Hunter Army Airfield in Savannah, GA in June 2007. In general helicopter engines exhaust particles size distributions observed at the engine nozzle and 4.14 m downstream locations showing the geometric mean diameters smaller than 50 nm for all engine power settings investigated in this study. For both locations, the geometric mean diameter increased as the engine power setting increased; this trend also holds true for the emitted particle number concentration. The growth of particle geometric mean diameter was found significant, 7 nm, only in the case of the idle power setting. Sulfur-to-sulfate conversion was found to be independent of the engine power setting. Emissions of both sulfur and sulfate increased as the engine power increased. When JP-8 fuel was used, particles smaller than 7 nm were found to increase in samples taken at the downstream location. The number concentration in this tail increased as the power setting increased. No such observation was found when FT fuel was used implying that the increased formation of nuclei particles in the plume downstream was likely to be caused by the sulfur and aromatic compounds in the JP-8 fuel. Total particulate carbon emissions increased as the engine power setting increased. Use of FT fuel reduced the elemental carbon emissions as it compared to the JP-8 fuel, and organic carbon emission at idle power but not at the higher powers. The reduction of elemental carbon by the FT fuel was attributed to the absence of aromatics (soot precursors) in the fuel. The OC/EC ratio was found to be in the range of 3-50 depending on the engine power setting. The aircraft emitted OC/EC was found to decrease as the engine power increased.