A requirement has been produced to perform a combination of experimental wind tunnel tests and numerical simulations to study the phenomena of acoustic refraction through the aerodynamic boundary layer over a fuselage. The aim of the research was to acquire and utilise experimental acoustic data to validate and develop numerical simulation methods in order to better understand the complex phenomena of noise transmission from Contra-rotating Open Rotor engines (CROR) into the aircraft cabin at cruise conditions.
The research utilised a significant number of existing test components and acoustic instrumentation to maximise the quantity of unique experimental data obtained. The tests were performed in the ARA 2.74m x 2.44m Transonic Wind Tunnel using an acoustic liner insert to create a near anechoic environment in the working section. An existing fuselage model was utilised, incorporating extensive steady and unsteady pressure sensor installations and a large array of microphones. An existing tonal noise source was also utilised.
Aircraft Research Association Ltd (ARA) proposed to apply its extensive experience and expertise in wind tunnel testing and model design and manufacture to lead this programme of work. ARA liaised with the CfP leader during the design and manufacture of the test hardware and subsequent wind tunnel tests, developing and exploiting current state of the art techniques including traversing mechanisms for the hot-wire probes and high-speed acoustic data recording.
The numerical acoustic simulations were performed by Free Field Technologies (FFT) using CAA (Computational Aero-Acoustics) codes, which were assessed and developed using the experimental acoustic data from the wind tunnel tests.
New aircraft concepts incorporating ultra-high bypass ratio turbofan or open rotor engine configurations offer significant increases in operational efficiency and economy with potential reductions in environmental impact due to reduced fuel burn. However, the open rotor configurations in particular pose an additional challenge, imposing relatively high noise levels at the surface of the aircraft fuselage, directly influencing passenger comfort and potentially requiring additional noise treatments with associated penalties of increased weight, fuel burn and environmental impact. The noise levels experienced at the fuselage surface can, however, be modified by acoustic refraction occurring during propagation through the boundary layer. In this context, the influence of turbulent boundary layer refraction on the propagation of noise to the surface of a three dimensional rear fuselage geometry had been investigated during the ENITEP project by the generation of an extensive, high speed experimental data base and comparison of these data with the results of computational aero-acoustic (CAA) simulations.
An existing, high speed wind tunnel model instrumented for the measurement of surface acoustics was modified to permit measurement of the steady and unsteady boundary layer aerodynamics. An in-flow noise source was employed to inject broadband and tonal noise at frequencies and separation distances representative of typical contra-rotating open rotor (CROR) configurations. In addition, the noise source was subject to an isolated, wind-on characterisation exercise. A qualitative insight into the refraction phenomena was gained by testing over a wide speed range up to a fuselage reference local Mach number of 0.78. Selected experimental test cases were simulated using Euler and RANS computational fluid dynamics (CFD) approaches to generate flow field inputs for the CAA tasks.
The noise source output was modelled numerically and the accuracy of existing CAA prediction tools was investigated by experimental/numerical comparisons of fuselage peak noise levels and directivities for the selected test cases. In addition, the steady refraction effect was estimated by comparison of CAA results obtained with and without the fuselage turbulent boundary layer mean flow represented.
CAA-derived peak noise levels generally exceeded those measured during test, although comparison of surface acoustic pressure distributions showed a generally good match in terms of directivity. Local acoustic attenuation by the fuselage boundary layer over the speed range M = 0.40 to 0.75 falling in the approximate range 1-12dB has been estimated from the CAA results. A dependence on Mach number and also noise source separation distance, particularly at low speed, has been observed.
In addition to continued development of existing acoustic prediction tools addressing the steady refraction phenomenon, the existing data set will also support future investigation of unsteady (scattering) refraction effects.