The aviation industry requires continuous reductions in greenhouse gas and noise emissions, which is resulting in a shift towards hybrid and fully electric aircraft. The propeller is an attractive form of propulsion system for electric aircraft as it can achieve high efficiencies and has excellent scalability potential. However, a major concern is the high community and cabin noise levels. It is possible to mitigate the propeller noise levels by reducing the blade loading and shifting the loading inboard, though this will often adversely impact the aerodynamic efficiency. Another method of reducing noise is by introducing blade sweep. This can result in destructive interference of sound waves. The downside is that sweep may impose significant structural deformations when the blade is loaded. These structural deformations may alter both the aerodynamic and acoustic performance. The objective of this thesis is to identify how these structural deformations will impact the aerodynamic and acoustic performance of low-noise, swept propeller blades. A parametric study was performed where the blade sweep was parametrised. The aerodynamic, acoustic and structural performances
of these propeller configurations were then analysed to identify the relations between these three disciplines.

A multidisciplinary propeller framework was used where aerodynamic, acoustic and structural tools could be coupled. For the aerodynamic analyses, a Vortex Lattice tool was developed, which can perform inviscid, incompressible simulations of any propeller geometry. The acoustics were calculated using Hanson’s Helicoidal Surface Theory, which models the blade sweep and dihedral as phase lag effects to model the interference between
sound waves. The structural tool used the Euler-Bernoulli beam theory for bending deformations and Saint-Venant’s torsion theory for twisting deflections. These three tools were coupled such that the final deformed shape of a loaded propeller could be obtained and evaluated. Each tool was validated individually using high-fidelity and experimental data, confirming that accurate results may be obtained within the tool’s applicable limitations.

The elastic blades were compared to their rigid counterparts to analyse how much the blade deformed, and how these deformations impacted the aerodynamic and acoustic performance. It was found that bending deformations are caused by the centrifugal forces, opposing any bending moments caused by the aerodynamic forces. The torsional deformations, mainly caused by the aerodynamic forces, result in a wash-out. Both the
bending deformations and wash-out of the blade reduce the local angle of attack, reducing the aerodynamic loads acting on the propeller blades. Since these structural deformations are concentrated near the blade tip, the local loading distributions are only impacted near the tip. This effectively shifts the loading inboard. The loss in performance can be significant if substantial sweep is applied, with losses of up to 20% of the thrust and torque. The efficiency, on the other hand, does not significantly change due to elasticity. The acoustic performance was mainly altered due to the changes in aerodynamic performance, with any changes in blade geometry having negligible effects. The overall sound pressure levels reduce by up to 2.5 dB, correlating with the reduction in thrust and torque. By eliminating the noise’s dependency on the overall thrust level, a clear correlation between the shifting of the local loading and the noise emissions was also found. These inboard shifting loads reduced the thrust-specific sound pressure by up to 1 dB.

It is possible to reduce the effect of elasticity by letting the structural forces impose a twisting moment, reducing the wash-out of the propeller blades. High bending deformations, or the addition of dihedral, allow the moment arm of the centrifugal loads to grow enough such that the twisting moments of the aerodynamic forces can be overcome. This can greatly diminish the aerodynamic performance loss. On the other hand, the associated increase of the peakiness of the local aerodynamic loading distributions may increase the noise
emissions even when a net loss in aerodynamic performance is still present.
Clear relations between the aerodynamic, acoustic, and structural performance of elastic, swept blades were identified. These changes in aerodynamics and acoustics are too great to be ignored for highly swept blades. Analyses of such flexible, highly swept blades clearly require the inclusion of structural calculations to accurately predict the aerodynamic and acoustic performance.

Boundary Layer Ingestion (BLI) is a promising technology for reducing the impact of aviation on the environment. By placing a propulsor on an aircraft such that it ingests the slower moving air within the boundary layer, a decrease in power consumption can be achieved. Evaluating the performance of BLI configurations remains a challenge though. Many conventional methods for computing the thrust and drag of an aircraft cannot be used due to the interference between the airframe and the propulsor. As a result, researchers have to rely on either low fidelity methods with limited accuracy, or on expensive CFD simulations. In this work, an attempt was made to create a computationally inexpensive method that can evaluate the drag, thrust and propulsive power of an axisymmetric fuselage with a BLI propulsor. The method proposed relies on potential flow modeling and makes use of a boundary layer model and several corrections.

Various potential flow methods with different assumptions are available to quantify the efficiency increase and thrust provided by a swirl recovery vane (SRV). In this paper, thrust coefficients and efficiency results obtained by different potential flow methods for the same SRV geometry at different advance ratios are presented. The methods include two VLM and four lifting line (LL) models with different assumptions. The models are compared in terms of accuracy with respect to RANS results and computational cost. This makes it possible to evaluate the benefits and drawbacks of neglecting or accounting for the presence of certain effects and modelling choices. The effects taken into account or deliberately neglected in different models include; finite propeller-SRV distance, nacelle presence, wake and free stream nonalignment, flow interaction between vane blades, the Kutta condition and SRV sweep. The wake angle behind the SRV is also varied and its effect on thrust coefficient is observed. In conclusion, accounting for the presence of a nacelle and finite slipstream distance respectively leads to 7.28% and 16.39% improvement in accuracy of the SRV thrust coefficient with negligible increase in CPU time. Not aligning the SRV wake with free stream direction has little impact on the computed thrust coefficient but causes the CPU time to increase steeply. Using a VLM based model rather a LL model and modelling vane interaction significantly increases CPU time whilst yielding the highest improvements in thrust coefficient accuracy (25.43% and 35.16%).

In recent years, there has been renewed interest in propeller research due to environmental concerns. Propellers are known for their low-speed efficiency and compatibility with electric motors. The utilization of regenerative energy during deceleration has been demonstrated to be beneficial for automobiles and has the potential to be applied to propeller-driven aircraft. This technology allows for the recharging of batteries and could increase the overall efficiency of the propulsion system by utilizing the regenerated energy in other flight maneuvers. The negative thrust mode of propellers has several additional advantages, including weight reduction, improved safety, and increased maneuverability. However, when operated in reverse thrust conditions, the performance of conventional propellers is suboptimal because they are designed for forward thrust. This suboptimal operation can lead to boundary layer separation on the blades, increasing broadband noise and potentially making it more dominant over tonal noise. To assess the relative importance between an isolated propeller's tonal and broadband noise sources under positive and negative thrust conditions, an experimental study was conducted using a scaled propeller setup in a low-turbulence wind tunnel (LTT). Aeroacoustic measurements were taken using a microphone array with 63 microphones placed in the floor and wall of the test section. The advance ratio was varied across different wind speeds at Reynolds numbers ranging from Re_c^0.7R = 0.8·10⁵ to Re_c^0.7R = 1.9·10⁵. Signal processing methods were used to analyze the data and quantify tonal and broadband noise sources, including fast Fourier transformation (FFT), phase averaging, beamforming, and blade element momentum (BEM) analysis. The results showed that the broadband noise component of the propeller significantly increased in negative thrust mode compared to positive thrust mode due to fully separated boundary layers. The tonal propeller components were often masked by interference from wind tunnel and motor noise, suggesting the need for further research to validate propellers in regenerative mode.

The angular momentum or swirl in the propeller slipstream is an energy loss. An effective method to recover the swirl and increase the propulsive efficiency is to use Swirl Recovery Vanes (SRVs). Former research showed an over-prediction in SRV thrust by a lifting line theory (LLT) model compared to wind tunnel experiments. The LLT model assumes an unbounded flow field from minus infinity to plus infinity, while in reality the SRV flow field is bounded due to a small axial spacing between the SRVs and the propeller plane. Using a 2D correction method, based on the application of the Kutta–Joukowski theorem in an axially confined domain combined with the thin airfoil approximation, a correction in the angle of attack can be computed for each blade element of the SRV. The LLT model including the correction method could match the SRV thrust of the experiments. Additionally, the original LLT thrust prediction was met with a corrected pitch angle of the SRVS. An SRV in off-design conditions would result in different correction angles. For a smaller thrust setting of the propeller, the thrust of the SRV will also be lower, hence the correction angle will be smaller. To make the correction method also applicable in off-design conditions, a fixed correction angle would be more practical since the whole vane can be rotated with a variable pitch to match the appropriate correction angle for the corresponding thrust setting. The fixed correction angle was determined by taking an average of all correction angles along the blade radius and the thrust prediction showed good agreement, with a negligible thrust loss of the SRV compared to correcting the pitch angle for each blade element separately. Finally, in the airfoil profile optimization, it was found that the profile drag of the profile is of less influence on the SRV thrust, hence more design freedom can be used to select the appropriate airfoil profile for the SRV.

With the increased focus on sustainability, aircraft are designed to reduce their emissions. One way to accomplish this is by increasing the wing aspect ratio, thereby increasing the aerodynamic efficiency, however this is not without consequences. Increased aspect ratio wings have a higher structural mass, are more susceptible to gust and maneuver loads and generally flutter at lower velocities. Due to the advent of both passive and active control techniques these issues can fortunately be solved by using gust load allevation (GLA), maneuver load alleviation (MLA) and flutter suppression.
Current aeroelastic testing facilities at Delft University of Technology include a gust generator and aeroelastic apparatus, used to suspend a passive wing section in the wind tunnel. The need for ad¬ ditional research on aeroelastic control in order to improve the sustainability and safety of aviation necessitates the development of a new wing section with aileron and spoiler control surfaces that is compatible with current facilities. The development, manufacturing and initial characterization and test¬ ing of this wing section is the subject of the present work.
As the new wing section includes a spoiler, a literature review is performed on this subject. Spoilers function by deflecting into the flow, causing separation aft of the spoiler and creating a large turbulent wake, resulting in a drastic decrease of lift. A linear potential flow model for spoiler aerodynamics developed by Brown and Parkinson was implemented in MATLAB with the intent of implementing this in future aeroelastic models. Verification of this model showed good agreements with original data presented in the paper describing the model.
The passive wing section was chosen as a basis for the new design. The position and size of the control surfaces are determined based on a review of experimental and operational applications. The new wing section was designed, resulting in a self¬contained model, including a single¬board computer, sensors and power supply. Actuation mechanisms were developed for the control surfaces, with a parametric device for control surface free play included in the aileron actuation mechanism. The new wing section was manufactured successfully and control software was implemented using Simulink.
A series of tests were performed to characterize the dynamic behavior of the wing section. Due to a combination of higher inertia and kinematics of the actuation mechanism, the usable bandwidth of the aileron is shown to be lower than that of the spoiler. Aerodynamic results show that the combined use of aileron and spoiler result in a reduction or reversal of the aerodynamic response of the wing. Gust load alleviation results with proportional control show an increase in damping by 1300% and a reduction in peak amplitude of 50% when using the spoiler. Results for the aileron are notice¬ ably less, with a decrease in amplitude of 15% and an increase of damping of 145%. The differences are attributed to both the differences in kinematics of the mechanisms as well as the greater absolute change in lift coefficient obtainable by the spoiler.

Due to the rising demand for short-range air travel and the desire for aircraft driven by electric propulsion, there has been a renewed interest in propeller research. Despite the high potential aerodynamic efficiency of propellers, their excessive noise emissions prevent a widespread use on aircraft. A gradient based optimization study is performed to assess the trade-off between aerodynamic and aeroacoustic performance of propellers. A blade-element-momentum theory(BEMT) approach with a dependence to the effects of blade sweep is used. The BEMT-model is combined with a frequency domain approach for tonal noise prediction. The optimization study shows that the advance ratio and pitch are important operational parameters during the climb phase to induce a noise reduction. The application of blade sweep proves useful as a design measure to reduce noise, given that the propeller operates in high speed conditions.

In the post processing of acquired wind tunnel force measurements, propeller-driven aircrafts need a preliminary correction of the propeller slipstream. Propeller slipstream effect has two major contributions to the flow field. The first one is the increase in dynamic pressure and the second is the addition of rotation in the flow field. Without a slipstream correction, standard wind tunnel corrections cannot be implemented because powering the propeller violates the underlying assumptions of the standard correction methods according to Eckert. As a result, Eckert has developed a thrust correction formula that only includes the increase of dynamic pressure. This research has been conducted to develop a novel correction to include both the increase in dynamic pressure and the effect of rotation induced by the propeller slipstream. Initially, a non-linear Surface Vorticity Panel Method (SVPM) was chosen in order to have a non-prescribed slipstream strength and shape. However, there were problems with convergence time and slipstream deformation. Therefore, the model was reduced to a linear model first suggested by Schroijen. The propeller was modeled as a mix of BEM and vortex theory. The wing and empennage was modeled by VLM. Fuselage was modeled after Multhopp’s vertical slit representation and a forced potential solution was implemented to simulate the wing root effect. The wind tunnel experiments were conducted in various angle of attacks and thrust ranges. Results show that the panel method can simulate the rotational nature of the slipstream accurately and the wing lift distribution is parallel to the literature. The new proposed correction was adequate at approximating the propeller slipstream lift and can produce closer results than the Eckert correction at some instances.

This thesis has two objectives: First, to investigate the behavior of typical twin engine commercial transport aircraft in ground effect under a wide range of conditions and geometry variations at the request of an aerospace company. And secondly, to add to the body of existing scientific knowledge on ground effect.In order to identify the opportunities for adding scientific knowledge, first a literature study was conducted. The conclusion of this study proved that there is insufficient research literature available on wing-bodies in the ground effect to answer the investigation launched by the aerospace company.A methodology was chosen to fulfill both objectives, namely generating ground effect performance data by means of CFD (Computational Fluid Dynamics) computations. The results are validated using existing wind tunnel measurements.It was determined which data points had to be calculated to cover the parameter space and it was established that the required effort could be completed within the time frame of this thesis. The landing high-lift configuration was the principal configuration to be investigated. Using this baseline, the impact of the following changes on ground effect performance were determined: a) changes to the high lift settings b) change to the engine thrust setting c) changes to landing gear deployment d) changes to the nacelle size e) changes to the wingtip geometry. After completing the calculation of all data points, it was also possible to develop a method that will enable engineers to calculate the effect of incremental changes in geometry and configuration, required for the initial stages of aircraft design.The main conclusions in order of importance are:For low to moderate angle of attack α, the ground effect is beneficial. Lift is increased by up to 3.5% of the total lift, drag is reduced by up to 33% of the total drag.However, as α>8.0°, the ground effect becomes less beneficial and the ground effect can even be detrimental, leading to a loss of lift of up to 5.5% while still having a drag benefit of up to 33%. The agreement between windtunnel data and CFD calculations is within 0.4% accuracy for the lift, and within 1.5% of the total drag value, except at the extreme case at moderate to high α and very close distance to the ground. Some of the differences can be explained in terms of practical problems in simulating the ground boundary conditions in the wind tunnel.The impact from a change to high lift setting and a change to thrust setting are most significant, with changes of up by up to 2% of total lift and up to 3.2% of total drag.There is a smaller but still quantifiable impact from the gear effect and the nacelle effect.A change to wing tip geometry has a minor impact on ground effect behavior, as lift is changed by up to 0.26% of total lift, only between 10.0°≤α≤12.0°.

The integration of the engine with the airframe is investigated in the German-Dutch Wind tunnels (DNW) using special scale engines called Turbofan Powered Simulators (TPS). The bookkeeping of thrust and drag must be clear. The TPS thrust is determined under static conditions in a calibration facility, and then subtracted in the wind tunnel test from the balance force of the aircraft model (airframe + TPS) to obtain the airframe drag including jet interference drag. A critical assumption valid for traditional turbofan engines is that the external flow does not affect the statically calibrated thrust. However, with the rise of more efficient engines with higher bypass ratio this may not be the case due to the close coupling between the engine flow and the wing. The objective of this Thesis is to identify the current limitations of the testing procedure as well as to produce scientific basis to deal with these limitations. This is achieved by means of a consistent thrust/drag bookkeeping combining numerical methods and a modified experimental setup.

External flow effects are identified by means of the analysis of theoretical models, the comparison with testing procedures undertaken in similar facilities, the review of standard bookkeeping techniques of full-scale turbofan engines and the analysis of previous test data. The change in thrust is quantified using a mathematical model integrated in an error propagation study by means of Monte Carlo simulations. The influence of the external flow and wing pressure field is further studied through a numerical analysis in RANS-SST for a very high bypass ratio TPS unit, and a Through-Flow Nacelle respectively. The latter configuration is also tested in the Low Speed Tunnel in DNW to investigate which instrumentation can be used to detect external flow effects in the future. In this case the velocity in the fan exhaust plane is measured with a hot wire and static pressure sensors are placed in the intake, exhaust and boattail.

The theoretical, numerical and experimental approach show that the external flow and wing pressure field change the conditions in the TPS exhaust with respect to static calibration. In the wind tunnel, the nozzle exhaust shear layer decreases in size as the difference in velocities between plume and free air is decreased, reducing the flow spreading rate and increasing the local pressure at the nozzle exit plane. The local Nozzle Pressure Ratio is reduced. This leads to flow suppression, the reduction of the fan mass flow and exhaust velocity. In addition, the scrubbing and boattail drag, currently bookkept as loss of thrust in the modified standard net thrust, are changed from static to wind tunnel conditions. These effects change the TPS thrust leading towards an improperly bookkeeping of the aircraft installation drag. The bias error produced by external flow effects is one order of magnitude higher than the random instrumentation error and should be corrected for, especially at low power settings. Differences decrease proportionally to the Fan Nozzle Pressure Ratio until chocked conditions are reached, where the freestream velocity has no influence in the TPS performance.

A possible solution lies in the advanced derivation of thrust and drag. The current approach neglects the thrust contribution from the nozzle exhaust to infinite downstream. According to the definition of the Jones thrust, a better solution lies in the assumption that the flow is expanded from the exhaust to infinity downstream without any transfer of energy of momentum. The decrease in mass flow and velocity can be effectively captured by pressure taps located at the intake or fan plane. A linear correlation exists between both stations, that can be used for the bookkeeping of the TPS thrust in the wind tunnel according to additional calibration in the wind tunnel. The new bookkeeping method can also be used to correct for the decrease in local jet exhaust Mach number from design conditions, the parameter of interest in engine/airframe integration tests.

The research presents and solves the limitations of the new generation of turbofan engines by accounting for the local conditions at the TPS exhaust due to external flow effects. The new thrust/drag bookkeeping method leads to optimized configurations by improving the accuracy of engine/airframe integration tests.

After the invention of the rocket mankind saw a new possibility to study the universe and other planets than Earth by space travel. Since then man has achieved to orbit Earth, land on the Moon and Mars. The constant increase in the demands of future space exploration mission make that also the equipment needs to be _t to the task. For example in 2020 NASA aims at returning a sample from Mars back to Earth [NASA Mars Exploration Program & Missions].

Junction flows occur when a boundary layer encounters an obstacle (such as a wing) on its path. They can occurr either in laminar or turbulent flow conditions but only the turbulent case is considered here. This type of flow is usually a 3D, anistoropic, unsteady, turbulent flow. These characteristics are the result of two main phenomena. The most important one is called ”horseshoe vortex”, it is due to separation of the flow ahead of the obstacle caused by the adverse pressure gradient generated by the obstacle. The second one can be referred as ”secondary corner vortex” and it is triggered by gradients of Reynolds stresses and thus has no counterpart in laminar flow. These types of secondary flow are responsible for an increase in aerodynamic drag due to the interaction of the approaching boundary layer and the boundary layer developing on the obstacle.

This report discusses the potential of pylon trailing edge blowing to reduce the adverse effects of airframe installation on the performance and noise emissions of pusher propellers. Both experimental and numerical analyses were performed, focusing on the pylon wake profiles, the propeller performance, and the propeller noise emissions. The experiments were executed in Delft University of Technology's Open Jet Facility using a scale model pylon and two propeller models (one powered and one windmilling). The numerical analysis combined an existing propeller lifting line code with analytic methods suited to predict the effects of installation on the propeller performance and noise emissions.

Nowadays, the evolution of multi-element airfoils presents a change of route, towards a "circular arch", in order to achieve the ultimate lift force [9]. Particularly, for wing-ap configurations, the enhancement of high lift performance, is of great interest and growing research at Delft University of Technology. Typically, the main approach with respect to delivering higher lift forces is by deecting the ap to higher angles of attack. However, the higher this angle is, the larger the curvature of the ow. As a result, separated ow dominates most of the ap's contour and significant part of the lift force might be lost. This problem is known to engineers for years and in order to overcome this lift loss, additional ow control techniques seem to be essential. Thus, with the goal of lift improvement in mind, the effect of drooped spoiler application was investigated experimentally, for a wing- Fowler ap model, at critical flight conditions i.e. high ap deection angle of 50o and Re = 1:6 _ 106 (40 m/s).