Low FPR Propulsion Noise and Performance in Ultra-Short Nacelles

Andreas Peters
Advisor: Prof. Spakovszky
Aircraft engines design trends tend towards higher bypass ratio, lower fan speed and fan pressure ratio (FPR) configurations for improved fuel burn, reduced emissions and noise. Low-pressure ratio fans offer increased propulsive efficiency and, besides enabling thermodynamic cycle changes for improved fuel efficiency, significant noise reductions can be achieved. As fan pressure ratios are reduced, innovative nacelle design concepts are required to limit the impact of larger diameter fans on nacelle weight and drag. Due to the shorter inlet ducts and the lower pressure ratios, fan design becomes more sensitive to inlet flow distortion at angle-of-attack or crosswind operating conditions and installation stagnation pressure losses. A second major consequence of short inlet and exhaust ducts is increased fan noise. Attenuation and shielding of blade-row interaction noise, fan broadband and BPF tone noise is limited in short nacelles. Since low FPR propulsors and their nacelles are more closely coupled than in current turbofan engines, inlet-fan and fan-exhaust nozzle interactions must be included in the aerodynamic and aero-acoustic assessment of the propulsion system. The goal of this effort is to define an advanced fan/nacelle design with benefits in performance, noise, and operability. Working towards this aim, the objectives are to (1) investigate inlet distortion transfer and determine the potential of endwall treatment and asymmetric geometries in short-nacelle designs using a coupled fan-inlet body force based approach, (2) interrogate flow features near the blade tip region and determine forces to improve performance, stability, and inlet distortion sensitivity, and (3) explore options to reduce fan source noise and radiated noise in short nacelles.

Pratt & Whitney low FPR, high-bypass ratio geared turbofan (source: AviationWeek.com, Jan. 2010)

Assessment of Propfan Propulsion Systems for Reduced Environmental Impact

Andreas Peters
Advisor: Prof. Spakovszky

Baseline CRP blade-tip vortex system: Front rotor tip-vortices and viscous wakes interacting with rear rotor contribute to interaction tone noise.

Current aircraft engine design studies tend towards higher bypass ratio, low-speed fan configurations in order to attain reductions in fuel consumption, emissions, and noise. Propfan (advanced turboprop) engine concepts investigated in the past by American, European, and Russian aircraft manufacturers have demonstrated significant benefits in these areas. However, considerable concern remains about the potential noise generated by propfan engines, including both inflight cabin noise and community noise during takeoff and approach. The overall goal of this project is to define an advanced CRP configuration with improved noise characteristics while maintaining the required aerodynamic performance for a given aircraft mission.
An aircraft performance, weight and balance, and mission analysis is conducted on a candidate CRP-powered aircraft configuration and a detailed aerodynamic design of a pusher CRP is carried out. Full wheel unsteady 3-D RANS simulations are then used to determine the time-varying blade surface pressures and unsteady flow features necessary to define the acoustic source terms.

Polar directivity at first interaction tone frequency: Implementing advanced source mitigation concepts in re-designed CRP significantly reduces interaction tone noise compared to baseline CRP design.

A frequency domain approach based on Goldstein’s formulation of the acoustic analogy for moving media and an existing single rotor noise method is extended to counter-rotating configurations. Using the developed CRP noise estimation method, the underlying noise mechanisms front-rotor wake interaction, aft-rotor upstream influence, hub-endwall secondary flows, and front-rotor tip-vortices to interaction tone noise are dissected and quantified. Based on this investigation, the CRP is re-designed for reduced noise incorporating a clipped rear-rotor and increased rotor-rotor spacing to reduce upstream influence, tip-vortex, and wake interaction effects. Maintaining the thrust and propulsive efficiency at takeoff, the noise is calculated for both designs. On the engine/aircraft system level, the re-designed CRP demonstrates significant noise reductions and the results suggest that advanced open rotor designs can possibly meet Stage 4 noise requirements.

Re-designed CRP for low noise: Clipping rear rotor reduces interaction of front rotor tip-vortices with rear rotor, thereby decreasing interaction tone noise.

Carbon Nanotube Bearing

Eugene Cook, Draper & MIT
Project lead at Draper: David J Carter (PI), Marc Weinberg, Peter Miraglia
Advisor: Prof. Spakovszky

Carbon nanotube rotor [not to scale]

Rotating Micro Electro-Mechanical Systems (MEMS) require rotary bearings, but current MEMS bearing technologies have drawbacks. Silicon rubbing on silicon wears out quickly. Gas bearings require a gas source, and are relatively low stiffness. A promising alternative, proposed and being pursued by the Charles Stark Draper Laboratory in collaboration with the GTL, is to use Carbon Nanotubes (CNTs). Multi-walled CNTs have a concentric-tube structure that lends itself to bearings. Each tube is strong, but there is little or no bonding between tubes, allowing them to slide relative to each other. However, the friction characteristics of these bearings are not precisely quantified. This project’s goal is to construct a simple CNT bearing rotary device, demonstrating MEMS and CNT compatible fabrication techniques, and allowing some data on the friction characteristics to be gathered. Applications of such a bearing technology could include microscale turbomachinery, as well as gyroscopes, pumps, and other rotating devices.

Propulsion System Integration and Noise Assessment of a Hybrid Wing-Body Aircraft

Dorian Colas, Dr. Elena De la Rosa Blanco, (former students: Leo Ng, Phil Weed)
Advisor: Prof. Spakovszky
Reducing the environmental impact of air travel is a major impetus to current research in aeronautics. A potential configuration that could enable step changes in fuel consumption, noise and emissions is a hybrid wing-body aircraft where a lifting fuselage is blended with the wings. Building on previous work from the Silent Aircraft Initiative, this project aims to develop a set of advanced predictive methods that will enable the design of a hybrid wing-body aircraft to meet NASA’s N+2 goals: (i) 25% less fuel burn, (ii) 80% less emissions, and (iii) 52 dB less noise compared to current aircrafts in service. MIT, in collaboration with Boeing, NASA, and UC Irvine, is defining the aircraft configuration and propulsion system to meet such goals.
One approach reducing propulsion system noise is to mount the engines above the airframe, utilizing the large planform area to shield the noise generated by the turbomachinery. A fast algorithm of medium-fidelity was developed based on Kirchoff’s diffraction theory to compute the shielding effect of the airframe using directivity compact sources. The method includes flight effects and is applicable to any kind of aircraft configuration.
An alternative configuration uses engines embedded in the airframe where the airframe boundary layer is ingested by a distributed propulsion system. In such configurations thrust and drag cannot be simply separated and instead the overall aircraft performance is assessed using a previously established power balance analysis. The design of an S-shaped inlet and distortion tolerant fan stage is also being pursued.
The approach is based on high-fidelity simulations of the coupled airframe, inlet and fan system using a body force based representation of the fan stage. Various design concepts will be explored with the goal to improve power savings and to mitigate inlet flow distortion and fan performance penalties.

A Noise Assessment Methodology for Highly-Integrated Propulsion Systems with Inlet Flow Distortion

Jeff Defoe
Advisor: Prof. Spakovszky
Reducing emissions, fuel burn, and noise are the main drivers for innovative aircraft design. Embedded propulsion systems, such as those used in hybrid wing-body aircraft, can offer fuel burn and noise reduction benefits but one of the major challenges in high-speed fan stages used in these embedded propulsion systems is inlet distortion noise, in particular multiple-pure-tone (MPT) noise. MPT noise consists of shaft-order tones due to shock interaction caused by small (+/- 0.2°) blade-to-blade stagger angle variations. The key challenge to MPT noise prediction in embedded propulsion systems is the inherent coupling of the acoustics and aerodynamics due to the presence of non-uniform flow. A new approach was developed to solve this problem based on a body force description of the fan blade row. The body force field not only represents the overall rotor characteristics, capable of capturing the distortion transfer effects, but for the first time is also used as the fan noise source. An unsteady perturbation force field captures the blade-to-blade flow variations that cause MPT noise.
The approach has been validated on NASA's Source Diagnostic Test fan and inlet, showing good agreement with experimental data for aerodynamic performance, acoustic source generation and far-field noise spectra. The approach was then employed with the objective of quantifying the effects of non-uniform flow on the generation and propagation of MPT noise. First-of-their-kind back-to-back coupled aero-acoustic computations were carried out, comparing the conventional inlet used in the validation case to a serpentine inlet. Both inlets delivered flow to the same NASA/GE R4 fan rotor at equal corrected mass flow rates. Although the source strength at the fan is increased by 45 dB in sound power level due to the non-uniform inflow, far-field noise for the serpentine inlet duct is increased on average by only 7 dB (3 dBA) overall sound pressure level in the forward arc. This is due to the redistribution of acoustic energy to frequencies below 11 times the shaft frequency and the apparent cut-off of tones at higher frequencies including blade-passing tones. The circumferential extent of the inlet swirl distortion at the fan was found to be 2 blade pitches, or 1/11th of the circumference, suggesting a relationship between the circumferential extent of the inlet distortion and the cut-off frequency perceived in the far field. The streamwise vortices associated with the inlet distortion locally alter the relative Mach number and create a region of evanescent wave behavior which is conjectured to be the cause of the changes in the far-field spectra.
In the final phase of the project, a parametric study of serpentine inlet designs is currently underway to quantify the effects of non-uniform flow on MPT noise generation and propagation. The results will be used in the formulation of a response surface model suitable for incorporation into NASA's ANOPP noise prediction framework. The understanding gained from the parametric study will also be useful in forming design guidelines for integrated propulsion systems.

The "Swirl Tube" - an Aircraft Drag Management Device to Reduce Noise and Fuel Burn

In collaboration with Dr. Parthiv Shah of ATA, and former Gas Turbine Laboratory student Hiten Mulchandani
Advisor: Prof. Zoltan Spakovszky
Aircraft on approach in high-drag and high-lift configuration create unsteady flow structures which inherently generate noise. For devices such as flaps, spoilers and the undercarriage there is a strong correlation between overall noise and drag such that, in the quest for quieter aircraft, one challenge is to generate drag at low noise levels.
The invention is a novel aircraft drag management concept to reduce aircraft noise during approach and to improve fuel burn in cruise. The idea is based on a swirling exhaust flow emanating, for example, from a jet engine nacelle (see figure) or a wing-tip mounted duct. A novel application is to exploit the low pressure in the vortex core of the swirling exhaust flow to generate drag. The idea is that in a steady streamwise vortex the centripetal acceleration of fluid particles is balanced by a radial pressure gradient. The very low pressure near the vortex core at the exit of the duct generates pressure drag. This streamwise vortex is in essence steady, yielding low noise levels and a quiet acoustic signature. To see a Quicktime movie of the swirl tube in action, click here (this is a large file so please be patient while it loads).

Loss Modeling of Turbine Tip Leakage Flows

Arthur Huang
Advisors: Prof. Greitzer, Dr. Tan

Formation of tip leakage vortex (Mischo, Behr, Abhari, 2008). DS1 is the dividing streamline between incidence-driven flow and pressure-driven flow. DS2 is the dividing streamline between flow ending up in the passage vortex and flow ending up in the leakage vortex.

A major source of inefficiency in a turbine results from pressure-driven flow leaking across the rotor tip from the pressure side to the suction side. The flow emerges from the tip gap in a jet, which rolls up into a vortex near the shroud/suction side corner of the blade passage. Entropy is generated as the leakage flow mixes with the mainstream flow. In addition to creating aerodynamic losses, tip leakage flows also transfer heat to the rotor tip so that an uncooled rotor tip may be damaged. Because of this, turbine designers introduce cooling flows, which bring with them their own mixing losses, as well as lower total work due to the cooling flow bypassing the combustor. This project aims to model the losses associated with turbine tip leakage in order to better design the rotor tip.

Schematic of tip leakage flow (Krishnababu et al., 2009)

Currently used models for aerodynamic tip leakage losses are correlations based on rotor tip lift coefficients, blade geometries, or simply an efficiency penalty proportional to the gap height. We have modeled the tip gap region as a series of 2D planes in the leakage streamline and radial directions. In each of these 2D planes, the flow is viewed as a 1D sudden expansion over a vena contracta. Mass and momentum control volume equations are solved to determine leakage mass flows and velocities, and hence entropy due to mixing, which is the efficiency loss. The next steps in this project are to conduct CFD analysis of tip leakage flow to determine whether the assumptions used in the modeling are reasonable and to develop and test models for losses from required cooling associated with the tip leakage.