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.

Compressor Aerodynamics in Large Industrial Gas Turbines for Powr Generation

Sitanun Sakulkaew
Advisor: Dr. Tan
The overall goal of the research is to improve the efficiency of large industrial gas turbines through improvement of compressor performance. Specifically, the research focuses on two important aspects of compressor science and technology. The first aspect addresses loss and flow blockage generation in high-speed multistage axial compressors to establish a design philosophy for high efficiency and for broadening the island of peak efficiency. The second aspect seeks to quantify the variation of efficiency as blade (rotor tip and stator hub) clearance approaches zero and its implication on peak efficiency in a multistage environment. The overall framework of the approach consists of using computational analyses to first establish the traceability of flow features as they impact compressor performance changes; this is then to be followed by experimental assessments.

A representative large industrial gas turbine for power generation.

Secondary Air Interactions with Main Flow in Axial Turbines

Metodi Zlatinov
Advisor: Dr. Tan
In the past decade, industrial gas turbines have by far become the most popular type of plant for power generation due to their compactness, low emissions and potential for power-heat cogeneration. In the effort to increase energy conversion efficiency, engineers have raised turbine inlet temperatures to well above the metal melting point. Turbine blades are generally protected by expensive thermal barrier coatings and various forms of internal and film cooling. However, in order to prevent hot gasses from being ingested into the unprotected cavities between rotating and stationary components, cool air bled from the compressor is used to purge the gaps at the endwalls. MIT, in collaboration with Siemens Energy and Siemens Corporate Research, is developing a computational approach to identify and understand loss generating flow processes of purge air interacting with mainstream flow in axial turbines.

Contours of change in volumetric entropy generation rate relative to a baseline case with no purge flow bring out the regions in a rotor blade passage that have modified losses as a consequence of purge flow injection from the hub gap upstream of the rotor. We have identified a number of effects that result in these changes: mixing out of the velocity difference between purge and mainstream flows, the generation of radial velocity gradients as a consequence of purge flow interacting with the passage vortex structures, and increased wetted and tip clearance flow losses due to a change of reaction. There is also a positive effect of reduced tip clearance losses when purge flow is injected from the shroud. These effects have been rigorously quantified, and their drivers have been pinpointed. This new knowledge provides clear guidelines for better turbine designs.

Ported Shroud Operation in Turbochargers

George Christou
Advisor: Dr. Tan
In recent years, due to environmental regulations, automotive turbochargers have been increasingly implemented to accomplish high powering and downsizing of internal combustion engines. The operability of the compressor is bound at low mass flow rate by the surge line. Surge is characterized as a breakdown of the flow with large pressure fluctuations that can cause rapid deterioration and in some cases failure of the compressor and the bearing system. A technique used to control the development of surge is by implementing a ported shroud at the inlet of the compressor. The ported shroud configuration is used to improve both the choke and surge lines on the compressor performance map.

Garret by Honeywell, Turbo Tech 103 (Expert), 2006

The overall goal of this research project, in collaboration with Honeywell Turbo Technologies, is to improve the performance of ported shroud centrifugal turbochargers. Specific goals include: providing an explanation of changes in the flow processes with and without ported shroud relative to compressor operation; identifying and quantifying loss mechanisms present in ported shroud centrifugal compressors; increasing the effective operating range by increasing surge and choke margins; and increasing the efficiency at off-design operating points.

Aeromechanic Response in High Performance Centrifugal Compressor Stage

Christopher Lusardi
Advisor: Dr. Tan
Impeller blades in centrifugal compressors are exposed to unsteady forces that can increase stress levels in the part, leading to premature structural failure. These unsteady forces may arise from different sources, a significant one of which is the unsteady pressure field from the impeller interaction with the downstream diffuser. The resulting time-varying loads can induce vibratory stresses in the blades that could be significantly higher than the steady-state stresses. There have been experimental/test/field observations of impeller blades breaking at trailing edge as well as leading edge that are traceable to unsteadiness associated with impeller-diffuser interactions. Furthermore, test data shows that not only does the unsteady impeller-diffuser interaction impact the impeller forced response characteristic but that it is also highly sensitive to impeller-diffuser gap variation; the situation with impeller blade leading edge exhibiting high response that is linked to impeller-diffuser interactions constitutes an upstream manifestation of a downstream stimulus. The overall goal is to characterize and identify the unsteady flow process in impeller-diffuser interaction on the observed impeller aeromechanic difficulty such as those described above.

Improved Performance Return Channel Design for Multistage Centrifugal Compressors

Anne-Raphaelle Aubry
Advisor: Prof. Greitzer
High-pressure multistage centrifugal compressors are used extensively in the energy industry across a wide variety of applications from refinery processes to gas injection for carbon capture and sequestration. Centrifugal compressor manufacturers are looking towards reduced radial and axial dimension compressors to meet customer’s demands for lower cost and higher reliability. As the dimensions of the centrifugal compressors shrink, the job of the return channel—which must turn the flow by 180° and remove the tangential component of the flow—becomes more difficult.
MIT, in collaboration with Mitsubishi Heavy Industries (MHI), is developing a novel return channel design for these multistage compressors with the objective of improving efficiency, while meeting geometry constraints.
Opportunities to improve “traditional” return channel design were identified in a previous investigation and qualitative best practices established. A quantitative assessment of these best practices is being undertaken, and use of an adjoint method to optimize the return channel shape is also under consideration. Candidate designs obtained with this adjoint method would then be refined to develop a design that addresses the desired performance improvements. Performance of the candidate design are to be assessed in a full-scale stage test at the MHI single-stage test facility.

Contours of radial velocity (left) and entropy production rate (right) in the baseline return vane show a region of reduced velocity flow on the vane suction side, surrounded by a region of high entropy production.

Small-Scale Gas Turbine Engines

Dr. Jürg Schiffmann
Advisor: Prof. Spakovszky
Small-scale gas turbine engines can provide much higher power densities than conventional batteries, and show promise as a portable and enduring power source. A 1kW class small-scale gas turbine generator is being developed for this purpose. Experimentally identified key issues for making the engine work are thermal management and rotordynamic stability.
The heat generated by the engine through windage losses in the bearings and in the generator needs to be removed and the rotor has to be effectively shielded from high temperature sources to ensure the mechanical integrity of the generator. The challenge is set by the small-scale architecture. The goal is to establish an appropriate thermal management scheme. The technical approach is based on a parametric thermal resistance network that is calibrated with experimental data.
The high rotational speeds required to reach acceptable aerodynamic efficiency call for gas lubricated bearings, which are known to have a low threshold of stability and are prone to large amplitude sub-synchronous vibration. Hence, another goal is to investigate these phenomena and develop design guidelines for high-speed gas bearings. Our approach is based on high level modelling of the bearings and to perform a sensitivity analysis to identify significant scaling effects.
Due to the small scale of these engines, the effect of the interactions between the different components on the efficiency is significant. Hence a future goal is to apply an integrated design and optimization approach to investigate alternative engine configurations.

A Methodology for Centrifugal Compressor Stability Prediction

Jonathan Everitt, Benny Kuschel, Arian Roeber, Dr. Jürg Schiffmann
Advisor: Prof. Spakovszky
Although centrifugal compressors exhibit the same type of instabilities as axial compressors, rotating stall and surge are characterized by a much broader spectrum of unstable behavior. The wide variety of instability behavior, along with the inherently complicated flow in such a machine, are primary reasons that rotating stall and surge in centrifugal compressors are less well understood than similar phenomena in axial compressors. As a consequence, a general theory or a criterion for the onset of instability in centrifugal compressors does not exist.
Instead, correlations are used to describe the surge point for a certain class of centrifugal compressors and to estimate the stability limit based on a priori knowledge of blade row characteristics. The major limitation of these methods is that these characteristics are only available after experimental measurements and thus the method is not of predictive nature. This research project is different from past efforts in that the prediction is purely based on centrifugal compressor geometry and does not rely on correlations or a priori knowledge of compressor characteristics.
The approach is two-pronged. Previous research indicates that for certain classes of centrifugal compressors the inception of instability is in the diffuser; however the underlying fluid mechanics is not well understood. To gain insight, unsteady 3-D RANS calculations are being carried out on the isolated diffuser using an inlet flow field derived from full stage calculations. The inlet conditions are perturbed with a short wavelength total pressure disturbance. If the disturbance grows in time, the machine is deemed dynamically unstable, and the operating point is determined to be the surge point. The inlet conditions can be varied to determine what flow features need to be present in order to trigger instability.
The second prong to the approach borrows ideas from previous work on axial compressors and consists of 3-D steady RANS calculations to determine the body force distributions representing the effects of discrete blades on the flow field. The body forces are then coupled to a 3-D unsteady Euler solver. The compressor model is then forced with a short wavelength body force impulse in the vaneless space. The goal is to demonstrate that the method can accurately predict both the stall point and the type of stall inception pattern (spike or modal waves) in centrifugal compressors.

Representation of bladed centrifugal compressor with axisymmetric body forces.

Unsteady pressure traces for pressure taps located at constant radius around vaneless space. Left: Response to forcing for an unstable operating point at 75% corrected design speed; two-lobed backward traveling modal pre-stall waves grow in amplitude. Right: Response to forcing for altered compressor at neutrally stable operating point at 70% corrected design speed; forward traveling, spike-like disturbance formed shortly after forcing.

Current Research at the Gas Turbine Laboratory

A Unified Approach for Vaned Diffuser Design in Advanced Centrifugal Compressors

Jonathan Everitt, Jürg Schiffmann
Advisor: Prof. Spakovszky
Modern internal combustion engines utilize high pressure ratio turbochargers combined with NOx control strategies to improve efficiency and reduce emissions. In such an application, the centrifugal compressor has to simultaneously achieve high efficiency, high pressure ratio and a broad operating range. To meet these requirements, the trend has been towards highly loaded compressors, utilizing high speed impellers with backward-leaning blades for extended operating range and vaned diffusers for enhanced pressure recovery with compact geometry.
The flow out of the impeller presents multiple challenges for the vaned diffuser: it is transonic, unsteady, and highly non-uniform in both axial and circumferential directions. The relative importance of each of these factors is not well understood: different researchers have reached different conclusions regarding, for example, the importance of the impeller outflow non-uniformity. Diffuser design therefore largely depends upon historical correlations, CFD simulation and careful experimentation.
The objective of the investigation is therefore to rigorously establish the links between diffuser geometry, performance, component matching and stability. The technical approach combines first principles based modeling with high-fidelity calculations and experiments using a unique swirling flow diffuser test rig at the Gas Turbine Laboratory. The goal is to develop design criteria and to define performance metrics expressed in terms of overall vane parameters and appropriately averaged inflow properties that can be applied in the preliminary design stage.

Early Gas Turbine History

1791 First patent for a gas turbine (John Barber, United Kingdom)
1904 Unsuccessful gas turbine project by Franz Stolze in Berlin (first axial compressor)
1906 GT by Armengaud Lemale in France (centrifugal compressor, no useful power)
1910 First GT featuring intermittent combustion (Holzwarth, 150 kW, constant volume combustion)
1923 First exhaust-gas turbocharger to increase the power of diesel engines
1939 World’s first gas turbine for power generation (Brown Boveri Company), Neuchâtel, Switzerland
(velox burner, aerodynamics by Stodola)

The man behind the early steam and gas turbine

Prof. Aurel Stodola (1859-1942)	Final Grade report at ETH Zurich: GPA = 6 (A+)  [courtesy ETH Zurich]

Seminal work by Stodola

English Translation of “Die Dampfturbinen” 1906	One-dimensional treatment – Velocity Triangles

World’s first industrial gas turbine – 1939

From the paper collection of Eddie Taylor, the first director of MIT GTL (1947–1969)

Commissioning of world’s first industrial gas turbine, Neuchatel, 1939 (Stodola at age 80) [picture courtesy ETH Zurich]	Drawing of first gas turbine [from Eddie Taylor’s paper collection]

World’s first industrial gas turbine “Neuchatel”, 2007

ASME Historic Mechanical Engineering Landmark in Birr, Switzerland (ALSTOM Headquarters)
GT Neuchatel operated for 63 years (generator failure in 2002) [picture courtesy of ASME]

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No future for aircraft gas turbine engines...

National Academy of Sciences, Committee on Gas Turbines (June 1940):
“In its present state … the gas turbine engine could hardly be considered a feasible application to airplanes mainly because of the difficulty in complying with stringent weight requirements imposed by aeronautics.”

But two people insisted and asked the right questions…

Frank Whittle and Hans von Ohain

First turbojet-powered aircraft – Ohain’s engine on He 178

The world’s first aircraft to fly purely on turbojet power, the Heinkel He 178.
Its first true flight was on 27 August, 1939.

First Patent by Whittle for Turbojet Engine in 1930

Whittle W-1 (1941)	Rolls-Royce Trent

“Whittle … stressed the great simplicity of his engine. Hives [Director of Rolls Royce]  commented, ‘We’ll soon design the bloody simplicity out of it.’ ” [From Genesis of the Jet]