This thesis focuses on designing and optimizing a shock table using Finite Element Analysis (FEA) to predict and assess Shock Response Spectrum (SRS) levels. The goal is to create a cost-effective,  and accurate shock table capable of replicating launch vehicle shock environments. A structured methodology is adopted, including concept generation, simulation, and design space exploration. The results show that careful tuning of impactor characteristics, fixture geometry, and material selection significantly improves shock table performance.

The Flared Folding Wingtip (FFWT) concept allows an aircraft’s wingspan to extend beyond airport gate limitations, while its freely moving wingtip can serve as a Gust Load Alleviation (GLA) device, potentially reducing the required structural weight of the wing.
To study the behavior of the FFWT in combination with a highly flexible wing, this project designed and manufactured a wind tunnel demonstrator, featuring ailerons on both the fixed wing section and the folding wingtip. The inboard aileron is sized to achieve a hypothetical roll rate of 15°/s, comparable to CS-25 certified aircraft, while the outboard aileron is designed to return the wingtip to its neutral position.
Sizing the wing spar of a highly flexible demonstrator while ensuring flutter occurs in the free hinge condition within the velocity range of the Open Jet Facility (OJF) at TU Delft proved to be a conflicting design requirement. This led to a predicted flutter speed of 20 m/s and a maximum wingtip deflection of 5.64%
A Ground Vibration Test (GVT) was conducted to validate the Finite Element Method (FEM) model. A comparison between the locked and free hinge conditions revealed that the locked hinge introduced unexpected flexibility into the structure, which suggests that the FEM representation should be refined to better capture the actual structural behavior.

Flared folding wing tips (FFWTs) improve aerodynamic efficiency but present aeroelastic challenges. This study develops an FFWT aeroelastic model using: (i) multibody constraint formulation for kinematics, (ii) non-linear static analysis for equilibrium under aerodynamic loading, and (iii) linearised dynamic analysis for time-dependent behaviour.

The multibody formulation defines the wing tip’s motion through hinge constraints, while the non-linear static analysis examines the effects of flare angles on equilibrium fold angles and reaction forces. The wing root bending moment (WRBM) decreases by 17% compared to a locked configuration but increases by 23.4% as the flare angle grows from 0o to 20o. For flare angles below 10o, the solver characteristics and initial equilibrium positions at lower velocities can lead to numerical issues such as zero-division errors and poorly conditioned matrices.

The linearised dynamic model, based on the static solution, is evaluated with different configurations: a locked hinge, a free hinge, and a locked-free hinge. Smaller flare angles allow higher fold angles but introduce minor anomalies in the inner wing tip’s response, while larger flare angles improve numerical stability yet cause more persistent oscillations. The locked-free case assesses hinge release during a gust encounter, where releasing the hinge at peak gust intensity leads to larger persistent oscillations. Artificial numerical diffusion and structural damping effectively reduce numerical noise in reaction moments, revealing underlying trends and improving stability.

This thesis investigates a novel load alleviation technique known as mini-tabs, small devices placed on the upper surface of wings to reduce lift. Reynolds-Averaged Navier-Stokes (RANS) simulations were conducted to analyze the impact of geometrical properties on mini-tab performance in reducing the load on a wing.

The initial simulations focused on a single mini-tab, varying parameters such as height, aspect ratio, chordwise, and spanwise positions. Results indicate that increasing height or aspect ratio leads to decreased lift due to flow deceleration at the leading edge and expansion of the separated flow region downstream of the mini-tab. This will also increase the pressure drag. Increasing the height or aspect ratio also accelerates the flow at the lower surface, further reducing the lift generated by the wing.

Placing mini-tabs closer to the wingtip exhibits similar effects on lift reduction as increasing the height of the mini-tabs. This is because the relative height of the mini-tab with the local chord length increases when the mini-tab is positioned closer to the wingtip. A maximum lift reduction is also achieved when the mini-tab is placed at a 60% chordwise position.

Subsequent simulations explored the performance of multiple mini-tabs, revealing that orienting them orthogonal to the wind direction may not be optimal due to gap formation between the mini-tabs that energizes the wake downstream of the mini-tabs and reduces the reduction in lift. Conversely, aligning the mini-tabs with the leading edge has a higher efficiency in the reduction of the lift.

Overall, the findings suggest that mini-tabs offer a viable method for reducing wing forces and for load alleviation

This thesis investigates the aeroelastic behaviour of a flexible wing with minitabs and their effectiveness in alleviating the peak gust loads. Minitabs, which are small deployable surfaces located on the upper surface of the wing, were examined through a series of static and dynamic Fluid-Structure Interaction (FSI) simulations, which evaluated the aerodynamic and structural responses when deployed and stowed. The transient response of deployment of the minitabs indicated a greater load reduction capability compared to their steady-state performance. Notably, the peak alleviation of the minitabs is achieved when the gust peak lies within the aerodynamic reaction time of the minitab. Evaluating the deployment of minitabs for various gust profiles further demonstrates that minitabs performed better in response to sharp, high-amplitude gusts, while their efficacy diminishes with longer gust lengths. When deployed at the ideal moment, the minitabs were found to reduce peak gust loads by an average of 5%. As a consequence of the mid-flight deployment and stowing of the minitabs, simulations revealed significant aeroelastic effects, characterized primarily by first bending mode oscillations at approximately 1 Hz. A sweep analysis of the deployment and stowing times revealed that delaying the actuation of the minitab beyond the ideal deployment or stowing time leads to an increased amplitude oscillations and extended recovery time to steady state conditions. The study evaluated the performance of the minitabs in response to various gust profiles characterized by different gust lengths. The minitabs demonstrated improved performance in alleviating the gust peak loads from short, sharp gusts. However, the effectiveness in reducing the peak loads diminished for gusts of longer gust length and subsequent longer recovery time. The study also identified gusts with frequencies close to the natural first bending frequency of the wing structure as critical cases, where resonant behaviour emerges, leading to large amplitude oscillations and prolonged recovery times. While minitabs effectively reduce peak lift and bending moments induced by gusts, they also introduce lower amplitude oscillations in the bending and torsion moments that persist over multiple cycles. While the peak moments do not exceed a magnitude that would raise a concern to the structural integrity, the subsequent prolonged multi-cycle oscillations in the moments raises concerns about potential fatigue life issues for the internal structure. In conclusion, while this research confirms the potential of minitabs as a viable load alleviation method, it emphasizes the need for further investigation. Comprehensive aeroelastic studies on various minitab configurations, geometries and placements over the wing will provide a complete understanding of the aeroelastic effects of the minitabs and determine an ideal configuration to maximize the load alleviation capability. Additionally, the original actuation method of the minitab involves the rotating them from a vortex generator position to the minitab position. This thesis simplified the actuation to a ON/OFF procedure. Further research is required to incorporate this motion to evaluate any aeroelastic effects resulting from this actuation motion.

This thesis presents the development, analysis, and testing of a comprehensive testing methodology for determining the mass, Center of Gravity (CoG), and principal Moments of Inertia (MoI) of small rocket stages. The methodology comprises five sequential tests, encompassing both static and dynamic procedures. The two static tests introduce innovative adaptations of established techniques. The multi-point weighing (MPW) method for the measurement of mass and horizontal CoG coordinates is modified into a suspended version, broadening its applicability, while the suspension method, measuring the CoG height, evolves into the Bifilar Suspension (BS) method, enhancing robustness and safety with the inclusion of an additional rope. The dynamic tests measure the body's MoI values, with the Bifilar Pendulum (BFP) method for roll MoI and the Compound Pendulum (CP) method for pitch and yaw MoI. Efficiency and cost-effectiveness are integral to the methodology, resulting in a streamlined testing campaign requiring minimal time and resource investment.

An analytical uncertainty analysis assessment explores the propagation of uncertainties in the employed methods, highlighting the impact of several parameters on the combined uncertainty on the results. Additionally, an analytical expression for the amplitude-dependent errors in dynamic tests is derived, providing a useful tool to predict such nonlinear effects. A simulation study numerically verifies the results from the uncertainty analysis, as well as the solution equations used for the methods.

The methodology's validation is carried out through three consecutive test campaigns. The results demonstrate the capability of the static tests to consistently determine mass and CoG coordinates with limited uncertainties. The BFP method achieves satisfactory accuracy, although unexpected deviations from the numerical predictions are observed. As for the CP method, multiple factors exert a large influence on the accuracy of the final results. Among the ones analyzed in this work are: the length of the ropes, the radius of gyration of the body, and the accuracy in the frequency measurement. Moreover, in both dynamic tests the type of suspension system is found to have an effect on the accuracy of the measurements.

While not all the intended objectives have been achieved, this thesis contributes to the understanding of testing methodologies for rocket stages, and offers insights into achieving accurate and precise results with simple and cost-effective methods.

The ever-growing aerospace industry's urgent need to reduce greenhouse gas emissions has ignited a surge of interest in hybrid or fully electric propulsion systems. The electrification of aircraft introduces the possibility for energy to be recovered during phases of flight where no power input is required, and previous research has demonstrated the potential for small energy balance enhancements using dual role propellers. The design and operation of dual-role propellers involves considering two opposing load cases: positive thrust and torque during propulsive operation, and negative thrust and torque in regenerative operation. Thus, the ideal blade shape for maximizing performance in each opposing operating condition will be very different. Structural tailoring of the composite blades may be an effective approach for reducing energy consumption during operation in both regimes. Accordingly, the primary objective of this research is to determine the extent of dual-role propeller performance improvements that may be obtained through the application of aeroelastic tailoring. Sensitivity studies were conducted to convey an understanding of how dual-role propeller performance is affected through variations in structural designs (ply orientations and thicknesses). Optimization studies were subsequently performed to identify the extent of performance enhancements yielded solely through aeroelastic tailoring for flexible constant- and variable-pitch propellers of fixed geometry, assuming installation on a reference aircraft, and evaluated over multiple cruise distances for a climb-cruise-descent mission.

The thesis objectives were achieved through the development and application of an aeroelastic analysis and optimization framework. Blade element momentum (BEM) theory was used for the aerodynamic model with engineering corrections for compressibility, effects of rotation, root- and tip-losses, and the turbulent wake state. Excellent agreement was obtained from comparisons with a previous BEM code, and reasonable agreement in performance trends was observed through comparisons with experimental data during verification and validation. A modified version of PROTEUS, an aeroelastic tailoring code that was developed at the TU Delft, was used for the structural model. The aerodynamic analysis routine of PROTEUS was modified to instead use the developed BEM code for the evaluation of loads, sensitivities, and performance. The structural model of PROTEUS was modified to feature centrifugal forces and different input structures that are more conducive to the analysis of rotor blades. The structural model implemented during this project accounts for geometric nonlinearities, as well as nonlinear loads through the application of a corotational framework, and it is capable of accurately representing the detailed 3D blade as a reduced-order Timoshenko beam element mesh through its use of a cross-sectional modeller. Finally, a tightly coupled aeroelastic analysis procedure was developed and applied, which ensures convergence through the minimization of a residual vector using Newton's method, and analytical sensitivities for all loads were included in the analysis. Excellent agreement was obtained during verification studies for both the structural and aeroelastic analyses.

Results from the optimization and sensitivity studies indicate that the flexible blades constructed out of symmetric-unbalanced laminates yield a significant variation in thrust and power through the presence of bend-twist and extension-shear coupling, which results in an increasing change in twist distribution with increasing deflection or elongation. Only small variations in performance were observed from symmetric-balanced laminates, as the minimal amount of coupling resulted in negligible twist deformations, which confirms that the presence of bend-twist and extension-shear coupling drives variations in performance obtained through aeroelastic tailoring. Furthermore, it was found that the presence of an aerodynamic wash-out effect augments the range of advance ratio values corresponding to high-efficiency operation during both propulsive and regenerative modes. An opposite trend was observed in the presence of a wash-in effect. Lastly, due to the significantly decreased loading in descent, combined with its small contribution towards the total mission energy consumption, effects of aeroelastic tailoring are significantly greater in propulsive conditions (climb and cruise) in comparison to regenerative conditions.

From optimization, it was found that the flexible constant-pitch propeller features an energy consumption that is between 0.7% and 1.5% lower than the energy consumption of the rigid propellers. Despite this consistent decrease in energy consumption, all optimal flexible constant-pitch propellers were found to regenerate between 3% and 25% less energy than the rigid variable-pitch propeller, and between 3% and 10% less energy than the rigid constant-pitch propeller. This further suggests that the application of aeroelastic tailoring is most-suitable for improving performance in propulsive mode, as the enhanced performance yielded in propulsive mode outweighs the degraded performance in descent by a significant enough margin to enable the flexible constant-pitch propeller to outperform all rigid propellers. Moreover, the flexible variable-pitch propeller naturally yielded an even better performance than the constant-pitch propeller, with a total mission energy consumption that is between 1.5% and 2.0% less than the energy consumption of the rigid propellers. It has thus been shown that aeroelastic tailoring can yield noticeable improvements in propeller performance, at least for the fixed mission profile and reference aircraft that was studied.

The flight tests which are part of the certification procedure that prove that no flutter occurs within the flight envelope are fraught with risk. To reduce the risk, numerous numerical analyses, wind tunnel and ground tests are performed, which together with the flight tests result in a long and costly test programme.

To reduce the risks associated with flutter flight testing, a new flutter test version of the Parametric Flutter Margin (PFM) method, specifically applied to wings undergoing large deflections is presented. The PFM method adds a stabilising parameter, such as a stabilising mass, to the model such that the flutter velocity is increased. By exciting the stabilising mass in one of the primary (i.e., x, y and z) directions while simultaneously measuring the response in these directions and repeating the excitation in other directions, the flutter margins that are associated with the original model can be determined. To demonstrate the method a wind tunnel test campaign was performed at TU Delft using the Delft Pazy Wing which can exhibit large nonlinear deflection. The wing was equipped with a flutter pod consisting of a shaker and stabilising mass that was placed at the mid-span position at the leading edge of the wing.

During the test campaign, three test series were performed. The first identified the flutter boundary through direct flutter tests, with the flutter onset and offset velocities being determined by actually hitting flutter that turned into an LCO. The second and third test series were the PFM measurements, where both SISO and MIMO PFM were applied to obtain the nominal flutter boundaries without actually hitting flutter. The different PFM results showed a maximum difference of 4.4 % between each other at an angle of attack, α, of 6°, with the difference found to be minimal at α = 0° and increasing with increasing angle of attack, which was as expected.

Compared to the directly measured flutter tests, the MIMO PFM results identified a flutter velocity of 4.8 % lower than the directly measured flutter offset velocity at an angle of attack of 4°, and the SISO PFM results identified the flutter velocity to be 8.2 % lower than the offset velocity at 4° angle of attack. The PFM-identified flutter frequencies showed a difference of less than 2 % compared to the direct flutter test, with the difference in identified flutter frequency between the SISO and MIMO PFM reaching a maximum of 2 %, which was also increasing with increasing angle of attack.

However, even with the successful application, several research areas remain open, and several points of improvement for the performed test campaign have been found. Although besides the points of improvement, the acquired data shows the potential of the PFM method for performing safer, shorter and consequently cheaper flight tests for the certification procedure of new aircraft configurations.

In the last few years, a new aircraft configuration has been developed: the electric Vertical Take-off/Landing aircraft (eVTOL). To improve aircraft efficiency, structures are lighter and more flexible. Due to the structure's flexible nature and the use of propellers for propulsion, eVTOL are prone to whirl flutter, which consists of a dynamic instability produced by the propeller aerodynamics and structure interaction.
This thesis aims to assess whirl flutter in a multirotor VTOL considering propellers in thrusting conditions. A semi-analytical model was developed and then implemented on MATLAB for two types of configurations: a conventional propeller and two counter-rotating propellers at the tip of a cantilever beam structure.
The propeller aerodynamics in thrust conditions were obtained through blade element momentum theory (BEMT) and extended to counter-rotating propellers. Then, the BEMT flow characteristics were coupled to a quasi-steady aerodynamics perturbation model using strip theory to obtain the forces produced by the propeller whirl motion. Additionally, the beam structure was modeled using a space frame finite element model, in combination with a two-degree of freedom system model for the propeller-pylon structure. Then, the structural system was coupled to the unsteady aerodynamic forces produced by the whirl motion obtaining the aeroelastic model from which the stability analysis was performed.
The model was used to analyze whirl flutter on a multirotor aircraft designed by the company Betronka SPA. It was found that the multirotor does not suffer from whirl flutter for the designed flight conditions, considering the conventional and counter-rotating configurations. Afterward, to study the thrust influence on whirl flutter, flow conditions and propeller angular speeds outside the designed limit were studied. The main findings are: the counter-rotating propeller configuration is more stable compared to the default configuration, increasing the whirl flutter speeds; thrust conditions stabilize the default configuration, increasing the propeller angular speed to encounter whirl flutter by 4%; thrust conditions are negligible for counter-rotating propeller configuration considering constant angular velocity.

The technical advancements achieved during the last decade have permitted the implementation of low-fidelity numerical methods to analyze simplified aeroelastic problems. However, the available computational power is still incapable of dealing with more realistic events involving complex structural dynamics and flow regimes. For this reason, experimental wind tunnel campaigns are still required to characterize relevant fluid-structure interaction events and to assist in the development of more accurate computational models. Nevertheless, conventional measurement systems present several constraints that limit the complete description of the phenomena being analyzed. This thesis is part of the research effort to define alternatives capable of providing full-field and simultaneous information of both flow and structural quantities in a non-intrusive way.

A wind tunnel experimental campaign has been developed in the Open Jet Facility to evaluate the ability of Lagrangian Particle Tracking (LPT) to characterize aerodynamic loads acting on large-scale flexible bodies. Three different load reconstruction approaches have been considered for this purpose: integral momentum conservation equation in its classical formulation, integral momentum conservation equation with an alternative formulation based on vorticity, and Kutta-Joukowski theorem. The accuracy and precision of each of the forenamed procedures have been assessed in both steady and unsteady inflow conditions taking force balance measurements as a reference.

The results of this thesis have proven the capabilities of LPT to determine the aerodynamic loads in aeroelastic wind tunnel investigations with acceptable accuracy and precision. Finally, it has also been demonstrated that the range of information provided by these techniques is superior to that achievable with conventional measurement devices, which improves the overall characterization of the fluid-structure interaction phenomena.

The flared folding wingtip (FFWT) is a concept presented by Airbus to increase the aerodynamic efficiency and provide a means of load alleviation. A gate-to-gate demonstration of this concept has already been presented with the AlbatrossONE, a scaled aircraft featuring the FFWT concept. In this demonstration, the aircraft fold the wingtips during taxiing to fulfil gate requirements and extends it before take-off for improved efficiency. In addition, they can be released during flight to provide load alleviation and reduce the roll rate reduction caused by the increase in the wingspan. The objective of this thesis is to further study the FFWT concept.

An aeroelastic wind tunnel experiment to identify the influence of the wing stiffness and hinge release threshold on the gust load alleviation performance of a folding wingtip design is presented in this study. Five models with different stiffness and tailoring properties are tested and the wing root bending moment at different conditions is compared to the response with locked hinge conditions to assess the impact on the gust load alleviation capabilities of the folding wingtip.

The results show that the structural properties do not have an important impact on the peak load alleviation but the hinge release threshold and timing do. Releasing with the correct timing can reduce significantly the peak loads. However, the dynamics of the system are affected by this release: the flutter speed is decreased and, although the performance can improve, load oscillations increase, which can be considered detrimental for reasons such as fatigue.

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.

This thesis embarked upon establishing and validating a design process to build and control a distributed TRIC concept seamless smart morphing wing to achieve simultaneous load alleviation, flutter suppression and drag minimization capabilities. A novel aeroelastic simulation tool was built to carry out composite wing skin optimization, with prototype testing carried out to design a flexible connection between modules. The work output was a fully built morphing wing which was successfully tested in the TU Delft Open Jet Facility. Control of the wing was partially made feasible through the development of a surrogate model of the system which was also validated with DIC testing.

Over the course of the last century, the aircraft-industry has continuously been developing and optimising the performance of their products. The rate of improvement has been decreasing, and it seems like an asymptote is being approached in the development of the current technologies. The demand for new innovative solutions is growing and these solutions shall allow the performance increase to bypass this stagnation. One of the solutions investigated is to develop an aircraft wing which has the capability to fly at optimal shape by autonomously optimising its own wing-shape during all phases of the mission. This could potentially lead to a severe increase in aerodynamic performance and controllability while actively minimising structural loads.