Simulated Forward Flight Research Articles

Measuring Surface Pressures on Rotor Blades Using Pressure-Sensitive Paint

This paper will present details of a pressure-sensitive paint system for measuring global surface pressures on rotor blades in simulated forward flight at the subsonic tunnel at the NASA Langley Research Center. The system was designed to use a pulsed laser as an excitation source and pressure-sensitive paint data were collected using the lifetime-based approach. The higher intensity of the laser allowed pressure-sensitive paint images to be acquired using a single laser pulse, resulting in a collection of images that can be used to determine blade pressure at a specific instant in time. This is extremely important in rotorcraft applications because the blades experience dramatically different flowfields depending on their position. In addition, there can be fluctuations on the blade that vary every cycle due to factors such as lead/lag, flapping, and twisting of the blade. These effects generally preclude the use of phase averaging and thus the need for acquiring the data in a single image pair. For this test, the entire upper surface of a blade was painted and imaged. After taking into account temperature effects on the pressure-sensitive paint, the results agree both qualitatively and quantitatively with both expected results as well as with pressure transducers. Several limitations of the technique have been identified and discussion of strategies to overcome them is also presented.

Blade Tip Pressure Measurements Using Pressure-Sensitive Paint

This paper discusses the application of pressure sensitive paint using laser-based excitation for measurement of the upper surface pressure distribution on the tips of rotor blades in hover and simulated forward flight. The testing was conducted in the Rotor Test Cell and the 14- by 22-ft Subsonic Tunnel at the NASA Langley Research Center on the General Rotor Model System (GRMS) test stand. The Mach-scaled rotor contained three chordwise rows of dynamic pressure transducers for comparison with PSP measurements. The rotor had an 11 ft 1 in. diameter, 5.45 in. main chord and a swept, tapered tip. Three thrust conditions were examined in hover, C(sub T) = 0.004, 0.006 and 0.008. In forward flight, an additional thrust condition, C(sub T) = 0.010 was also examined. All four thrust conditions in forward flight were conducted at an advance ratio of 0.35.

The flow around a model helicopter main rotor in ground effect

Wind tunnel measurements of the wake below and ahead of a model helicopter main rotor in simulated forward flight in ground effect are presented. The wind tunnel used was equipped with a rolling road, and the ground speed was matched to the wind tunnel speed for a representative simulation in the wind tunnel of forward flight over the ground. Particle image velocimetry was used to investigate the structure of the wake, and it was observed that the moving ground had a remarkable effect on the flow; the wake is closer to the rotor and its size is reduced compared with the stationary ground case. The detailed distribution of vorticity within the wake is affected by the moving ground, and the mechanism for this is discussed in the paper.

Effects of Forward Flight on Jet Mixing Noise from Fine-Scale Turbulence

It is known experimentally that a jet in forward flight radiates less noise than the samejetin a static environment. At a forward flight Mach number of 0.2, the noise reduction, depending on the jet operating conditions, could be as large as 4-5 dB in the sideline directions. In the past, a way to predict flight effects was to use the method of relative velocity exponent. Another method was to extrapolate measured static jet noise to the flight condition by means of scaling formulas. Both methods are semi-empirical. The fine scale turbulence jet mixing noise theory of Tam and Auriault (Tam, C. K. W., and Auriault, L., Jet Mixing Noise from Fine Scale Turbulence, AIAA Journal, Vol. 37, No. 2, 1999, pp. 145-153) is extended for application to jets in simulated forward flight. It will be shown that the calculated noise spectra at different simulated forward flight Mach numbers for both supersonic and subsonic jets compare well with experiments. The effects of forward flight on the sources of fine-scale turbulent jet mixing noise is also investigated. It is found that in the presence of forward flight the dominant noise sources move downstream. The turbulence intensity and the size of turbulent eddies responsible for noise emission are reduced.

Effects of forward flight on jet mixing noise from fine-scale turbulence

It is known experimentally that a jet in forward flight radiates less noise than the samejetin a static environment. At a forward flight Mach number of 0.2, the noise reduction, depending on the jet operating conditions, could be as large as 4-5 dB in the sideline directions. In the past, a way to predict flight effects was to use the method of relative velocity exponent. Another method was to extrapolate measured static jet noise to the flight condition by means of scaling formulas. Both methods are semi-empirical. The fine scale turbulence jet mixing noise theory of Tam and Auriault (Tam, C. K. W., and Auriault, L., Jet Mixing Noise from Fine Scale Turbulence, AIAA Journal, Vol. 37, No. 2, 1999, pp. 145-153) is extended for application to jets in simulated forward flight. It will be shown that the calculated noise spectra at different simulated forward flight Mach numbers for both supersonic and subsonic jets compare well with experiments. The effects of forward flight on the sources of fine-scale turbulent jet mixing noise is also investigated. It is found that in the presence of forward flight the dominant noise sources move downstream. The turbulence intensity and the size of turbulent eddies responsible for noise emission are reduced.

Supersonic nozzle mixer ejector

The intent of this article is to describe recent experimental findings relative to supersonic nozzle mixer ejector performance. Such ejectors are a candidate means to reduce jet noise of commercial supersonic aircraft during takeoff and landing. The mixer ejector concept involves the introduction of an array of large-scale, low-intensity streamwise vortices into the downstream mixing duct, which enhances mixing through an inviscid stirring process. This results in increased ejector pumping performance and more completely mixed flows exiting the ejector shroud. Past experimental and analytical investigations of mixer ejectors have been confined to low-speed subsonic flows, and low primary temperatures (less than 2000°F). In this flow regime, ejector static pumping benefits of over 100% were achieved relative to conventional ejector designs. The goal of the present study was to evaluate mixer ejector performance in the high-temperature, supersonic primary flow regime. A convergentdivergent primary lobed nozzle (i.e., mixer nozzle) was designed and tested at choked pressure ratios in an ejector. Ejector pumping and exit plane mixing were measured for the mixer ejector and a conventional slot nozzle ejector. The two configurations were operated at a nozzle exit Mach number of 1.5 (nozzle pressure ratio = 3.4), a primary fluid total temperature of 1000°F, and a simulated forward flight Mach number of 0.1. Results indicate that properly designed lobed nozzles can increase supersonic ejector pumping by over 15%, relative to conventional slot primary nozzles.

Relationship Between Static, Flight, and Simulated Flight Jet Noise Measurements

Jet noise data are acquired in three different forms: static, flight, and simulated forward flight. Often data must be transformed from one type to one of the other two. A number of studies exist defining the relationship between these three types of measurements. These studies are deficient because they do not explicitly consider the jet noise source characteristics or do not clearly distinguish between mean square pressure, power spectral density, and one-third octave band sound pressure level measurements. In the current study, the relationships between static, subsonic flight, and subsonic simulated forward flight one-third octave band sound pressure level measurements were defined.

Noise of a model helicopter rotor due to ingestion of isotropic turbulence

A theoretical and experimental investigation of the noise of a model helicopter rotor due to the ingestion of grid-generated, isotropic turbulence is described. Simulated forward flight and vertical ascent tests were performed with a 0·76 m diameter, articulated model rotor in the United Technologies Research Center Acoustic Research Tunnel. Far field noise spectra and directivity were measured in addition to inflow turbulence intensities, length scales and spectra. Measured inflow turbulence statistics and rotor operating parameters were employed in a theoretical procedure to predict turbulence ingestion noise spectra and directivity. This theoretical formulation represented an absolute level prediction method in that empirical or adjustable constants were not employed. From this study it is concluded that incident turbulence represents a potentially important source of rotor narrowband random (quasi-tonal) and broadband noise. With due regard to the absolute level nature of the prediction method, agreement between theory and experiment is reasonable: quasi-tonal noise at low frequency tends to be overpredicted whereas mid-frequency tonal noise and high frequency broadband noise are well predicted. This theory, therefore, provides a means to predict the contribution of turbulence ingestion noise to overall rotor noise spectra and directivity in cases where reasonable estimates of incident turbulence statistics can be made. Ingestion of main rotor wake turbulence by tail rotors in flight and ingestion of atmospheric turbulence by main and tail rotors in hover would be expected to be the most important sources of turbulence ingestion noise in full scale applications.

Noise and detectability characteristics of small-scale remotely piloted vehicle propellers

Several small-scale propeller configurations applicable to a conceptual remotely piloted vehicle (RPV) aircraft were designed, fabricated, and tested to determine their performance, acoustic, and detectability characteristics. The tests were conducted in static and simulated forward flight conditions in a wind tunnel. Propellers tested included tractor, pusher, and ducted configurations. The acoustic data obtained were used to determine the slant range and altitude of no detection of each propeller configuration. The acoustic and detectability characteristics of small-scale RPV propellers were found to be significantly different from those of the large-scale propellers. An increase in forward velocity resulted in a significant drop in the SPL's at higher harmonics of blade passage frequency. As expected, tip speed had a very strong effect on noise and detectability in forward flight. It was found that most of the propellers tested would be detected at one of the first few harmonics (mostly first or second) of their blade passage frequency. Of the propellers tested at the design values of thrust and tip speed, three-bladed propellers were generally less detectable than either twoor four-bladed propellers for most of the forward velocities considered. Pusher and ducted propellers were found to be more detectable than their tractor and open counterparts.

Aeroelastic stability of complete rotors with application to a teetering rotor in forward flight

The derivation of a set of non-linear coupled flap-lag-torsion equations of motion for moderately large deflections of an elastic, two-bladed teetering helicopter rotor in forward flight is concisely outlined. The following degrees of freedom are included in the mathematical model: rigid body flapping, rigid body lead-lag, elastic bending in flap and lead-lag, blade root torsion, and shaft torsion. Quasi-steady aerodynamic loads are considered and the effects of reversed flow are included. The aeroelastic stability of the complete rotor is investigated by using a linearized system of equations of motion. The equilibrium position about which the equations are linearized is obtained by considering the trim state of the helicopter, in true or simulated forward flight conditions. The sensitivity of the aeroelastic stability boundaries to interblade structural and mechanical coupling is illustrated by comparing the complete rotor stability boundaries with those obtained from a single blade analysis for a number of hover and forward flight cases.

Effect of Simulated Forward Flight on Subsonic Jet Exhaust Noise

A wind-tunnel model program was conducted to investigate the effect of aircraft forward speed on the jet noise characteristics of a single subsonic jet. The tests were performed in the United Technologies Research Center Acoustic Research Tunnel, a low-noise open-jet anechoic wind tunnel designed specifically for acoustic testing. Electrically heated air was supplied to a convergent nozzle located on the tunnel centerline. The nozzle diameter of 2.4 in. produced a tunnel to nozzle area ratio of 220. Microphones were located at a radius of 50 nozzle diameters in the anechoic chamber outside the tunnel flow. Nozzle operating conditions covered a range of velocities up to 1670 fps, and temperatures to 900°F. Wind-tunnel speeds ranged from near zero to 350 fps. Measured noise data were corrected for tunnel shear layer refraction and moving medium convection effects, thus simulating the effects on the jet noise sources of relative velocity during an aircraft flyover. Measured overall sound pressure level reductions due to relative velocity were correlated with relative velocity raised to an exponent that increased towards the aft angles. The results showed general agreement with recent National Gas Turbine Establishment (NGTE) test results obtained in a large wind tunnel.

Electromechanical Simulation of Helicopter Blade Responses to Random Excitation During Forward Flight

The response of a model helicopter rotor blade to random excitation while in simulated forward flight is studied analytically and experimentally by means of an electromechanical apparatus. Generalized transfer functions are defined which relate steady-state responses in bending, flapping, and torsion modes to a sine input. Responses occur at the input and side-band frequencies. These transfer functions are then used along with excitation power spectra to predict the nonstationary time-averaged power spectrum of the response. Validity of the transfer function analysis is investigated by means of the electromechanical model which includes analog computer simulation of the interaction of blade deflections and aerodynamic load. Generalized transfer functions are measured for sinusoidal excitation. They are then used with measured excitation power spectra to predict the response, and the result is compared with measured response power spectra. Agreement is generally good for low advance ratio, but discrepancies diverge with increasing advance ratio.

The Use of a Runway Vehicle for Testing Lifting Rotors in Simulated Forward Flight

During recent years a major research programme has been undertaken at the National Gas Turbine Establishment into the subject of circulation control and its application to lifting rotors. In the case of rotors intended to be used on helicopters and stowed rotor aircraft, a very important part of the research work is the examination of rotor behaviour in forward flight. Traditionally, this has been investigated mainly on full-scale aircraft but more recently model scale experiments have been done in wind tunnels, particularly in the USA where there are a number of suitably large tunnels.