R. Vos
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17 records found
1
Morphing Trailing Edge Wings for Transport Aircraft
A conceptual assessment
Aviation was responsible for 2.5% of global energy-related CO2 emissions in 2023, and the market is projected to double over the next two decades. This growth conflicts with international targets to reach net-zero emissions by 2050. Of the available mitigation pathways, reducing in-flight energy consumption is the most immediately deployable, as it does not depend on new ground infrastructure. Aerodynamic drag is the main contributor to in-flight energy demand and is largely governed by the main wing. Since fixed wings are optimized for a single design point, they operate off-design during much of a mission. Morphing wings can adapt their shape in flight, offering a route to lower drag across a wider operating envelope.
This thesis assesses the effect of implementing chordwise two-degree-of-freedom (2-DoF) compliant camber-twist morphing trailing-edge wings on the sizing, configuration, and performance of CS-25 transport aircraft. Block fuel consumption is used as the primary performance metric, as it captures the trade-off between improved aerodynamic efficiency and the sizing penalties introduced by morphing systems.
Three methodological contributions support this assessment. First, a parameterization method was developed to describe the morphed shape of an arbitrary base airfoil using the rear spar location and two pseudo-deflection angles. Control points on the camberline are rotated and fitted with a fourth-order polynomial, after which the airfoil surfaces are reconstructed while preserving suction-side arc length to reflect compliant structure constraints. Applied at multiple spanwise stations, this method represents camber-twist morphing at wing level.
Second, the Aircraft Trimmed Performance Analysis Tool (ATPAT) was developed to evaluate trimmed cruise aerodynamic efficiency of morphing-wing aircraft. ATPAT combines a vortex lattice method with strip theory and sweep theory to resolve induced, viscous, pressure, and wave drag with runtimes of seconds to one minute per evaluation. The tool was validated against experimental and computational data on the Fokker 100 and integrated into the Initiator aircraft design toolbox with Bayesian and gradient-based optimizers.
Third, two uncertain sizing inputs were explicitly addressed: morphing system mass and achievable high-lift increment. Morphing system mass was treated as a parameter bounded by the specific mass of Fowler flaps, while RANS CFD estimated a conservative sectional lift increase of 0.63 at full deflection.
The methodology was applied to the ATR 72-600 and Fokker 100. For both aircraft, morphing designs were heavier and required larger wing areas than conventional designs due to high-lift limitations, making field performance the main constraint. Relaxing landing field length by about 20% brought sizing outcomes closer to the reference aircraft.
For the ATR 72-600, aerodynamic gains did not offset the penalties. The morphing design matched conventional block fuel only on the harmonic mission and used 6% and 4% more fuel on shorter and maximum-fuel missions. For the Fokker 100, a block-fuel reduction of up to 3.6% was achieved on the harmonic mission after a 17.5% relaxation of landing field length, with sizing nearly identical to the reference aircraft. Its transonic cruise also offers additional drag reduction potential through wave drag reduction.
The results remain subject to uncertainty. The relative contributions of sizing effects and wave drag reduction were not isolated, morphing system mass remains uncertain, and high-lift performance estimates are based on limited airfoil data. More detailed experimental validation and segmented mission analysis are recommended to refine the estimated benefits of morphing wings.
...
This thesis assesses the effect of implementing chordwise two-degree-of-freedom (2-DoF) compliant camber-twist morphing trailing-edge wings on the sizing, configuration, and performance of CS-25 transport aircraft. Block fuel consumption is used as the primary performance metric, as it captures the trade-off between improved aerodynamic efficiency and the sizing penalties introduced by morphing systems.
Three methodological contributions support this assessment. First, a parameterization method was developed to describe the morphed shape of an arbitrary base airfoil using the rear spar location and two pseudo-deflection angles. Control points on the camberline are rotated and fitted with a fourth-order polynomial, after which the airfoil surfaces are reconstructed while preserving suction-side arc length to reflect compliant structure constraints. Applied at multiple spanwise stations, this method represents camber-twist morphing at wing level.
Second, the Aircraft Trimmed Performance Analysis Tool (ATPAT) was developed to evaluate trimmed cruise aerodynamic efficiency of morphing-wing aircraft. ATPAT combines a vortex lattice method with strip theory and sweep theory to resolve induced, viscous, pressure, and wave drag with runtimes of seconds to one minute per evaluation. The tool was validated against experimental and computational data on the Fokker 100 and integrated into the Initiator aircraft design toolbox with Bayesian and gradient-based optimizers.
Third, two uncertain sizing inputs were explicitly addressed: morphing system mass and achievable high-lift increment. Morphing system mass was treated as a parameter bounded by the specific mass of Fowler flaps, while RANS CFD estimated a conservative sectional lift increase of 0.63 at full deflection.
The methodology was applied to the ATR 72-600 and Fokker 100. For both aircraft, morphing designs were heavier and required larger wing areas than conventional designs due to high-lift limitations, making field performance the main constraint. Relaxing landing field length by about 20% brought sizing outcomes closer to the reference aircraft.
For the ATR 72-600, aerodynamic gains did not offset the penalties. The morphing design matched conventional block fuel only on the harmonic mission and used 6% and 4% more fuel on shorter and maximum-fuel missions. For the Fokker 100, a block-fuel reduction of up to 3.6% was achieved on the harmonic mission after a 17.5% relaxation of landing field length, with sizing nearly identical to the reference aircraft. Its transonic cruise also offers additional drag reduction potential through wave drag reduction.
The results remain subject to uncertainty. The relative contributions of sizing effects and wave drag reduction were not isolated, morphing system mass remains uncertain, and high-lift performance estimates are based on limited airfoil data. More detailed experimental validation and segmented mission analysis are recommended to refine the estimated benefits of morphing wings.
...
Aviation was responsible for 2.5% of global energy-related CO2 emissions in 2023, and the market is projected to double over the next two decades. This growth conflicts with international targets to reach net-zero emissions by 2050. Of the available mitigation pathways, reducing in-flight energy consumption is the most immediately deployable, as it does not depend on new ground infrastructure. Aerodynamic drag is the main contributor to in-flight energy demand and is largely governed by the main wing. Since fixed wings are optimized for a single design point, they operate off-design during much of a mission. Morphing wings can adapt their shape in flight, offering a route to lower drag across a wider operating envelope.
This thesis assesses the effect of implementing chordwise two-degree-of-freedom (2-DoF) compliant camber-twist morphing trailing-edge wings on the sizing, configuration, and performance of CS-25 transport aircraft. Block fuel consumption is used as the primary performance metric, as it captures the trade-off between improved aerodynamic efficiency and the sizing penalties introduced by morphing systems.
Three methodological contributions support this assessment. First, a parameterization method was developed to describe the morphed shape of an arbitrary base airfoil using the rear spar location and two pseudo-deflection angles. Control points on the camberline are rotated and fitted with a fourth-order polynomial, after which the airfoil surfaces are reconstructed while preserving suction-side arc length to reflect compliant structure constraints. Applied at multiple spanwise stations, this method represents camber-twist morphing at wing level.
Second, the Aircraft Trimmed Performance Analysis Tool (ATPAT) was developed to evaluate trimmed cruise aerodynamic efficiency of morphing-wing aircraft. ATPAT combines a vortex lattice method with strip theory and sweep theory to resolve induced, viscous, pressure, and wave drag with runtimes of seconds to one minute per evaluation. The tool was validated against experimental and computational data on the Fokker 100 and integrated into the Initiator aircraft design toolbox with Bayesian and gradient-based optimizers.
Third, two uncertain sizing inputs were explicitly addressed: morphing system mass and achievable high-lift increment. Morphing system mass was treated as a parameter bounded by the specific mass of Fowler flaps, while RANS CFD estimated a conservative sectional lift increase of 0.63 at full deflection.
The methodology was applied to the ATR 72-600 and Fokker 100. For both aircraft, morphing designs were heavier and required larger wing areas than conventional designs due to high-lift limitations, making field performance the main constraint. Relaxing landing field length by about 20% brought sizing outcomes closer to the reference aircraft.
For the ATR 72-600, aerodynamic gains did not offset the penalties. The morphing design matched conventional block fuel only on the harmonic mission and used 6% and 4% more fuel on shorter and maximum-fuel missions. For the Fokker 100, a block-fuel reduction of up to 3.6% was achieved on the harmonic mission after a 17.5% relaxation of landing field length, with sizing nearly identical to the reference aircraft. Its transonic cruise also offers additional drag reduction potential through wave drag reduction.
The results remain subject to uncertainty. The relative contributions of sizing effects and wave drag reduction were not isolated, morphing system mass remains uncertain, and high-lift performance estimates are based on limited airfoil data. More detailed experimental validation and segmented mission analysis are recommended to refine the estimated benefits of morphing wings.
This thesis assesses the effect of implementing chordwise two-degree-of-freedom (2-DoF) compliant camber-twist morphing trailing-edge wings on the sizing, configuration, and performance of CS-25 transport aircraft. Block fuel consumption is used as the primary performance metric, as it captures the trade-off between improved aerodynamic efficiency and the sizing penalties introduced by morphing systems.
Three methodological contributions support this assessment. First, a parameterization method was developed to describe the morphed shape of an arbitrary base airfoil using the rear spar location and two pseudo-deflection angles. Control points on the camberline are rotated and fitted with a fourth-order polynomial, after which the airfoil surfaces are reconstructed while preserving suction-side arc length to reflect compliant structure constraints. Applied at multiple spanwise stations, this method represents camber-twist morphing at wing level.
Second, the Aircraft Trimmed Performance Analysis Tool (ATPAT) was developed to evaluate trimmed cruise aerodynamic efficiency of morphing-wing aircraft. ATPAT combines a vortex lattice method with strip theory and sweep theory to resolve induced, viscous, pressure, and wave drag with runtimes of seconds to one minute per evaluation. The tool was validated against experimental and computational data on the Fokker 100 and integrated into the Initiator aircraft design toolbox with Bayesian and gradient-based optimizers.
Third, two uncertain sizing inputs were explicitly addressed: morphing system mass and achievable high-lift increment. Morphing system mass was treated as a parameter bounded by the specific mass of Fowler flaps, while RANS CFD estimated a conservative sectional lift increase of 0.63 at full deflection.
The methodology was applied to the ATR 72-600 and Fokker 100. For both aircraft, morphing designs were heavier and required larger wing areas than conventional designs due to high-lift limitations, making field performance the main constraint. Relaxing landing field length by about 20% brought sizing outcomes closer to the reference aircraft.
For the ATR 72-600, aerodynamic gains did not offset the penalties. The morphing design matched conventional block fuel only on the harmonic mission and used 6% and 4% more fuel on shorter and maximum-fuel missions. For the Fokker 100, a block-fuel reduction of up to 3.6% was achieved on the harmonic mission after a 17.5% relaxation of landing field length, with sizing nearly identical to the reference aircraft. Its transonic cruise also offers additional drag reduction potential through wave drag reduction.
The results remain subject to uncertainty. The relative contributions of sizing effects and wave drag reduction were not isolated, morphing system mass remains uncertain, and high-lift performance estimates are based on limited airfoil data. More detailed experimental validation and segmented mission analysis are recommended to refine the estimated benefits of morphing wings.
Aviation faces increasing societal and regulatory pressure to reduce environmental impact, particularly noise in densely populated areas. Electric propulsion and Urban Air Mobility concepts aim to enable quieter flight. However, propeller generated aerodynamic noise remains a dominant contributor to the overall acoustic signature. Modelling broadband noise, associated with turbulent inflow and blade surface turbulence, is essential for realistic noise assessment.
This thesis develops a framework to evaluate broadband noise prediction models for (electrically driven) propellers in hover and forward flight. Two model are assessed: the classical Brooks, Pope, and Marcolini (BPM) model and the data-driven Gill and Lee (GL) model. Both are verified against literature and validated and compared using scaled hover experiments and full-scale flyover measurements.
Results show that the GL model performs well in hover but overpredicts noise in forward flight. The BPM model provides more consistent predictions across operating conditions. ...
This thesis develops a framework to evaluate broadband noise prediction models for (electrically driven) propellers in hover and forward flight. Two model are assessed: the classical Brooks, Pope, and Marcolini (BPM) model and the data-driven Gill and Lee (GL) model. Both are verified against literature and validated and compared using scaled hover experiments and full-scale flyover measurements.
Results show that the GL model performs well in hover but overpredicts noise in forward flight. The BPM model provides more consistent predictions across operating conditions. ...
Aviation faces increasing societal and regulatory pressure to reduce environmental impact, particularly noise in densely populated areas. Electric propulsion and Urban Air Mobility concepts aim to enable quieter flight. However, propeller generated aerodynamic noise remains a dominant contributor to the overall acoustic signature. Modelling broadband noise, associated with turbulent inflow and blade surface turbulence, is essential for realistic noise assessment.
This thesis develops a framework to evaluate broadband noise prediction models for (electrically driven) propellers in hover and forward flight. Two model are assessed: the classical Brooks, Pope, and Marcolini (BPM) model and the data-driven Gill and Lee (GL) model. Both are verified against literature and validated and compared using scaled hover experiments and full-scale flyover measurements.
Results show that the GL model performs well in hover but overpredicts noise in forward flight. The BPM model provides more consistent predictions across operating conditions.
This thesis develops a framework to evaluate broadband noise prediction models for (electrically driven) propellers in hover and forward flight. Two model are assessed: the classical Brooks, Pope, and Marcolini (BPM) model and the data-driven Gill and Lee (GL) model. Both are verified against literature and validated and compared using scaled hover experiments and full-scale flyover measurements.
Results show that the GL model performs well in hover but overpredicts noise in forward flight. The BPM model provides more consistent predictions across operating conditions.
Liquid Hydrogen Tank Layouts in Blended Wing Body Aircraft
Conceptual Design and Comparative Evaluation
Hydrogen-fueled aircraft promise to achieve significantly reduced climate impact compared to their kerosene counterparts. However, while hydrogen has a significantly higher gravimetric density than kerosene, its volumetric density is much lower, meaning that even in a liquid state, it takes up approximately four times the volume. This results in significant challenges in integrating hydrogen storage in aircraft. It has been suggested that blended wing body (BWB) aircraft may provide a better-suited alternative for hydrogen integration compared to the conventional tube-and-wing (TAW) configuration.
Prior research has investigated hydrogen-fueled BWB aircraft, highlighting different potential tank integration strategies. To the author's knowledge, no direct comparison of these hydrogen configurations highlights their relative performance impact compared to a consistent kerosene baseline aircraft. Therefore, this work compares different hydrogen tank configurations identified in prior research to a kerosene BWB baseline under consistent top-level aircraft requirements (TLARs), including a design range of 2500 nmi and 239 passengers. Technology assumptions for a 2050 entry into service (EIS) are applied. The analysis is complemented through the development of hydrogen and kerosene TAW configurations that fulfill the same TLARs.
The comparison is performed in three steps. First, a BWB is designed for each tank configuration using conceptual design methods. Second, the four tank layouts are compared and the layout is identified that imposes the lowest energy penalty on the BWB. Third, the integration penalty of the hydrogen BWB is compared to that of the hydrogen TAW.
A BWB employing a combined tank configuration, with hydrogen tanks located beside and aft of the passenger cabin, is found to experience the lowest integration penalty. The BWB experiences a 12.5% penalty in block energy for the design mission, compared to a penalty of 10.5% experienced by the TAW. This indicates that the BWB is less well suited to hydrogen integration than the TAW under the specified TLARs. In addition, sensitivity studies are conducted to evaluate the impact of integration assumptions on the BWB. These show that reasonable variations in the assumptions do not change the conclusion of this study. The findings do not eliminate hydrogen-fueled BWB aircraft as a viable alternative to hydrogen-fueled TAW designs. In fact, literature shows that BWB configurations still offer an inherent efficiency advantage, although their higher integration penalty must be considered in the overall trade-off. ...
Prior research has investigated hydrogen-fueled BWB aircraft, highlighting different potential tank integration strategies. To the author's knowledge, no direct comparison of these hydrogen configurations highlights their relative performance impact compared to a consistent kerosene baseline aircraft. Therefore, this work compares different hydrogen tank configurations identified in prior research to a kerosene BWB baseline under consistent top-level aircraft requirements (TLARs), including a design range of 2500 nmi and 239 passengers. Technology assumptions for a 2050 entry into service (EIS) are applied. The analysis is complemented through the development of hydrogen and kerosene TAW configurations that fulfill the same TLARs.
The comparison is performed in three steps. First, a BWB is designed for each tank configuration using conceptual design methods. Second, the four tank layouts are compared and the layout is identified that imposes the lowest energy penalty on the BWB. Third, the integration penalty of the hydrogen BWB is compared to that of the hydrogen TAW.
A BWB employing a combined tank configuration, with hydrogen tanks located beside and aft of the passenger cabin, is found to experience the lowest integration penalty. The BWB experiences a 12.5% penalty in block energy for the design mission, compared to a penalty of 10.5% experienced by the TAW. This indicates that the BWB is less well suited to hydrogen integration than the TAW under the specified TLARs. In addition, sensitivity studies are conducted to evaluate the impact of integration assumptions on the BWB. These show that reasonable variations in the assumptions do not change the conclusion of this study. The findings do not eliminate hydrogen-fueled BWB aircraft as a viable alternative to hydrogen-fueled TAW designs. In fact, literature shows that BWB configurations still offer an inherent efficiency advantage, although their higher integration penalty must be considered in the overall trade-off. ...
Hydrogen-fueled aircraft promise to achieve significantly reduced climate impact compared to their kerosene counterparts. However, while hydrogen has a significantly higher gravimetric density than kerosene, its volumetric density is much lower, meaning that even in a liquid state, it takes up approximately four times the volume. This results in significant challenges in integrating hydrogen storage in aircraft. It has been suggested that blended wing body (BWB) aircraft may provide a better-suited alternative for hydrogen integration compared to the conventional tube-and-wing (TAW) configuration.
Prior research has investigated hydrogen-fueled BWB aircraft, highlighting different potential tank integration strategies. To the author's knowledge, no direct comparison of these hydrogen configurations highlights their relative performance impact compared to a consistent kerosene baseline aircraft. Therefore, this work compares different hydrogen tank configurations identified in prior research to a kerosene BWB baseline under consistent top-level aircraft requirements (TLARs), including a design range of 2500 nmi and 239 passengers. Technology assumptions for a 2050 entry into service (EIS) are applied. The analysis is complemented through the development of hydrogen and kerosene TAW configurations that fulfill the same TLARs.
The comparison is performed in three steps. First, a BWB is designed for each tank configuration using conceptual design methods. Second, the four tank layouts are compared and the layout is identified that imposes the lowest energy penalty on the BWB. Third, the integration penalty of the hydrogen BWB is compared to that of the hydrogen TAW.
A BWB employing a combined tank configuration, with hydrogen tanks located beside and aft of the passenger cabin, is found to experience the lowest integration penalty. The BWB experiences a 12.5% penalty in block energy for the design mission, compared to a penalty of 10.5% experienced by the TAW. This indicates that the BWB is less well suited to hydrogen integration than the TAW under the specified TLARs. In addition, sensitivity studies are conducted to evaluate the impact of integration assumptions on the BWB. These show that reasonable variations in the assumptions do not change the conclusion of this study. The findings do not eliminate hydrogen-fueled BWB aircraft as a viable alternative to hydrogen-fueled TAW designs. In fact, literature shows that BWB configurations still offer an inherent efficiency advantage, although their higher integration penalty must be considered in the overall trade-off.
Prior research has investigated hydrogen-fueled BWB aircraft, highlighting different potential tank integration strategies. To the author's knowledge, no direct comparison of these hydrogen configurations highlights their relative performance impact compared to a consistent kerosene baseline aircraft. Therefore, this work compares different hydrogen tank configurations identified in prior research to a kerosene BWB baseline under consistent top-level aircraft requirements (TLARs), including a design range of 2500 nmi and 239 passengers. Technology assumptions for a 2050 entry into service (EIS) are applied. The analysis is complemented through the development of hydrogen and kerosene TAW configurations that fulfill the same TLARs.
The comparison is performed in three steps. First, a BWB is designed for each tank configuration using conceptual design methods. Second, the four tank layouts are compared and the layout is identified that imposes the lowest energy penalty on the BWB. Third, the integration penalty of the hydrogen BWB is compared to that of the hydrogen TAW.
A BWB employing a combined tank configuration, with hydrogen tanks located beside and aft of the passenger cabin, is found to experience the lowest integration penalty. The BWB experiences a 12.5% penalty in block energy for the design mission, compared to a penalty of 10.5% experienced by the TAW. This indicates that the BWB is less well suited to hydrogen integration than the TAW under the specified TLARs. In addition, sensitivity studies are conducted to evaluate the impact of integration assumptions on the BWB. These show that reasonable variations in the assumptions do not change the conclusion of this study. The findings do not eliminate hydrogen-fueled BWB aircraft as a viable alternative to hydrogen-fueled TAW designs. In fact, literature shows that BWB configurations still offer an inherent efficiency advantage, although their higher integration penalty must be considered in the overall trade-off.
APPU Aircraft Empennage Design
Multidisciplinary Design Optimization of a Cruciform or T-tail Empennage
The APPU project aims to lower the threshold of installing hydrogen-driven and boundary-layer-ingesting propulsion systems in the short term. This is done by replacing the Auxiliary Power Unit with an Auxiliary Power & Propulsion Unit (APPU). To study these concepts, the Airbus A321neo is taken as a reference baseline. The implementation of the APPU system on the A321neo aircraft necessitates a redesign of the existing empennage. The research objective of this thesis project is to investigate how an optimal empennage design for an aircraft equipped with an APPU system differs from an optimal empennage design for an aircraft without such a system. Multidisciplinary design optimization is used to minimize fuel weight for different aircraft configurations by optimizing the empennage geometry. Four different disciplines are identified. A weight discipline estimates the empennage weight based on the empirical Raymer equations. Two separate aerodynamic disciplines are implemented with different fidelity levels. The low-fidelity version is based on AVL's vortex lattice method, expanded with a constant friction coefficient drag model to account for viscous effects. The high-fidelity version is based on FlightStream's panel method. FlightStream proved to be infeasible for use in this study due to long run times and limited mesh robustness. Therefore, only the low-fidelity version is used to generate the final results. The static stability and control of the aircraft are ensured through a set of constraints that require specific stability and control derivatives to remain within defined limits. These derivatives are provided by the stability and control discipline that is based on AVL as well. The performance discipline applies the Breguet equation to convert the aircraft weight and aerodynamic performance into an estimate of the fuel load required to complete the design range. ParaPy is used as a multi-model-generator to provide the required input geometries for the disciplines. The study reveals that optimal empennage designs for aircraft equipped with an APPU system differ notably from optimal empennage designs for conventional aircraft. Both configurations benefit from high aspect ratios and reduced tailplane areas. However, APPU-equipped aircraft require a low cruciform tail to accommodate the hydrogen tank and reduced sweep angles to position the aerodynamic center aft without intersecting the propulsor plane. The low-fidelity aerodynamics discipline is unable to optimize the airfoils. AVL’s modeling of only the camber line makes it unsuitable for optimizing symmetric profiles, such as those on vertical tailplanes. Furthermore, the horizontal tailplane airfoils showed limited variation from the initial design. This is likely due to the dual-parameter definition of the camber line, as the class-shape transformation parameterization method is applied independently to both the upper and lower surfaces of the airfoils.
...
The APPU project aims to lower the threshold of installing hydrogen-driven and boundary-layer-ingesting propulsion systems in the short term. This is done by replacing the Auxiliary Power Unit with an Auxiliary Power & Propulsion Unit (APPU). To study these concepts, the Airbus A321neo is taken as a reference baseline. The implementation of the APPU system on the A321neo aircraft necessitates a redesign of the existing empennage. The research objective of this thesis project is to investigate how an optimal empennage design for an aircraft equipped with an APPU system differs from an optimal empennage design for an aircraft without such a system. Multidisciplinary design optimization is used to minimize fuel weight for different aircraft configurations by optimizing the empennage geometry. Four different disciplines are identified. A weight discipline estimates the empennage weight based on the empirical Raymer equations. Two separate aerodynamic disciplines are implemented with different fidelity levels. The low-fidelity version is based on AVL's vortex lattice method, expanded with a constant friction coefficient drag model to account for viscous effects. The high-fidelity version is based on FlightStream's panel method. FlightStream proved to be infeasible for use in this study due to long run times and limited mesh robustness. Therefore, only the low-fidelity version is used to generate the final results. The static stability and control of the aircraft are ensured through a set of constraints that require specific stability and control derivatives to remain within defined limits. These derivatives are provided by the stability and control discipline that is based on AVL as well. The performance discipline applies the Breguet equation to convert the aircraft weight and aerodynamic performance into an estimate of the fuel load required to complete the design range. ParaPy is used as a multi-model-generator to provide the required input geometries for the disciplines. The study reveals that optimal empennage designs for aircraft equipped with an APPU system differ notably from optimal empennage designs for conventional aircraft. Both configurations benefit from high aspect ratios and reduced tailplane areas. However, APPU-equipped aircraft require a low cruciform tail to accommodate the hydrogen tank and reduced sweep angles to position the aerodynamic center aft without intersecting the propulsor plane. The low-fidelity aerodynamics discipline is unable to optimize the airfoils. AVL’s modeling of only the camber line makes it unsuitable for optimizing symmetric profiles, such as those on vertical tailplanes. Furthermore, the horizontal tailplane airfoils showed limited variation from the initial design. This is likely due to the dual-parameter definition of the camber line, as the class-shape transformation parameterization method is applied independently to both the upper and lower surfaces of the airfoils.
Master thesis
(2025)
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O. van Voorst, M.F.M. Hoogreef, A. Heidebrecht, S. Giovani Pereira Castro, R. Vos
The increasing urgency to mitigate climate change has underscored the need to transition from conventional fossil-based aviation fuels, such as kerosene, to sustainable alternatives. Hydrogen stands out due to its potential to significantly reduce greenhouse gas emissions, making it a promising energy carrier for the aviation sector. However, adopting hydrogen presents substantial challenges, with the development of specialized fuel containment systems being one of the foremost obstacles. Double-walled tanks employing vacuum insulation offer an effective solution for cryogenic hydrogen storage, but they require a robust supporting structure for the inner vessel. Various solutions have been proposed to support the inner vessel of cryogenic tanks; however, despite the diversity of designs, there remains a noticeable lack of comprehensive research focusing specifically on the structural behaviour and feasibility of these inner vessel supporting structures for commercial aircraft applications. This study proposes a novel fibre-based suspension technique for the inner vessel of a double-walled integral tank designed for liquid hydrogen storage in large commercial aircraft. A finite element model was developed to evaluate the structural interaction between the inner and outer vessels and the supporting fibres, enabling structural sizing optimization to assess the impact of added loads. A parametric study was conducted to explore the influence of fibre design parameters on structural performance, mass, displacement, and thermal behaviour. Key design guidelines were established. First, using more than two longitudinal anchoring points results in the unwanted transfer of bending loads from the outer to the inner vessel. Second, while increasing the number of circumferential fibres reduces peak loads and displacements, the associated anchoring mass is the primary limiting factor, as thermal conduction was found to be negligible. Lastly, fibre orientation should prioritize low stiffness in the contraction direction to minimize tensile forces under initial filling. This should be combined with fibres angled in the longitudinal direction to improve longitudinal stiffness and displacement control. The results confirm the structural feasibility of the suspension system, showing only a marginal structural mass increase of approximately 1.88 % compared to a baseline integral tank without internal support. These findings provide practical guidance for the implementation of fibre-based suspension systems in cryogenic tank structures, supporting the development of hydrogen storage solutions for aviation.
...
The increasing urgency to mitigate climate change has underscored the need to transition from conventional fossil-based aviation fuels, such as kerosene, to sustainable alternatives. Hydrogen stands out due to its potential to significantly reduce greenhouse gas emissions, making it a promising energy carrier for the aviation sector. However, adopting hydrogen presents substantial challenges, with the development of specialized fuel containment systems being one of the foremost obstacles. Double-walled tanks employing vacuum insulation offer an effective solution for cryogenic hydrogen storage, but they require a robust supporting structure for the inner vessel. Various solutions have been proposed to support the inner vessel of cryogenic tanks; however, despite the diversity of designs, there remains a noticeable lack of comprehensive research focusing specifically on the structural behaviour and feasibility of these inner vessel supporting structures for commercial aircraft applications. This study proposes a novel fibre-based suspension technique for the inner vessel of a double-walled integral tank designed for liquid hydrogen storage in large commercial aircraft. A finite element model was developed to evaluate the structural interaction between the inner and outer vessels and the supporting fibres, enabling structural sizing optimization to assess the impact of added loads. A parametric study was conducted to explore the influence of fibre design parameters on structural performance, mass, displacement, and thermal behaviour. Key design guidelines were established. First, using more than two longitudinal anchoring points results in the unwanted transfer of bending loads from the outer to the inner vessel. Second, while increasing the number of circumferential fibres reduces peak loads and displacements, the associated anchoring mass is the primary limiting factor, as thermal conduction was found to be negligible. Lastly, fibre orientation should prioritize low stiffness in the contraction direction to minimize tensile forces under initial filling. This should be combined with fibres angled in the longitudinal direction to improve longitudinal stiffness and displacement control. The results confirm the structural feasibility of the suspension system, showing only a marginal structural mass increase of approximately 1.88 % compared to a baseline integral tank without internal support. These findings provide practical guidance for the implementation of fibre-based suspension systems in cryogenic tank structures, supporting the development of hydrogen storage solutions for aviation.
Operational effectiveness of hydrofoils for littoral craft
Investigating the potential of hydrofoils for high-speed daughter craft in amphibious operations
This research explores the use of submerged hydrofoil systems for landing craft, with the goal of evaluating its operational effectiveness. Using a systems engineering approach, various submerged hydrofoil configurations and propulsion systems are evaluated. A design method is developed which uses iterative parametric modelling and assessment of the craft's dynamic equilibrium across its full speed range. The resulting designs are then compared to a planing benchmark vessel. The findings show a resistance reduction of 35% at cruise speeds, leading to a 50% increase in range. Seakeeping analysis of one submerged hydrofoil design also demonstrated that hydrofoils can increase the sustained speed in waves and effectively reduce peak vertical accelerations. The results highlight that significant gains can be made with respect to range, sustained speed, and safety, which exemplifies the potential of hydrofoil system integration despite added design complexity.
...
This research explores the use of submerged hydrofoil systems for landing craft, with the goal of evaluating its operational effectiveness. Using a systems engineering approach, various submerged hydrofoil configurations and propulsion systems are evaluated. A design method is developed which uses iterative parametric modelling and assessment of the craft's dynamic equilibrium across its full speed range. The resulting designs are then compared to a planing benchmark vessel. The findings show a resistance reduction of 35% at cruise speeds, leading to a 50% increase in range. Seakeeping analysis of one submerged hydrofoil design also demonstrated that hydrofoils can increase the sustained speed in waves and effectively reduce peak vertical accelerations. The results highlight that significant gains can be made with respect to range, sustained speed, and safety, which exemplifies the potential of hydrofoil system integration despite added design complexity.
Flying-V Modeling and Robust Controller Design
Design of a C* Longitudinal Controller within the H-infinity Signal-Based Mixed Sensitivity Framework
As the aviation industry faces growing environmental and societal pressures, novel concepts such as the Flying-V emerge. Nevertheless, its unconventional configuration poses unique challenges in terms of stability and control, highlighting the pressing need for advanced control systems development. A gap in the literature arises concerning the lack of system robustness in the presence of uncertainties and the absence of robust controllers designed for the aircraft. Hence, the current study aims to increase the maturity of the flight control system of the Flying-V concept by implementing a C* longitudinal controller, while guaranteeing robustness stability and performance against uncertainty, adequate performance in the presence of disturbances and measurement noise, and compliance with Level 1 handling qualities. The Flying-V model is implemented in MATLAB/Simulink which served as the foundation for the controllers synthesis. Three designs are developed, which include a continuous time, a modified continuous time, and a modified continuous time multi-modeling controllers. These are conducted within the standard and signal-based H-infinity mixed sensitivity frameworks. Extensive analysis of the controllers is performed in terms of stability assessments, linear and nonlinear time domain simulations, and uncertainty sensitivity. Results demonstrate that the feedback of a combined signal demand such as the C* parameter provides, for a single input multiple-output system, a balanced disturbance rejection at the plant outputs. Conclusions are drawn in terms of the feedback controller structure, highlighting that high gain is necessary at low frequencies for disturbance rejection and roll-off at high frequencies allows control signal reduction and measurement noise attenuation. Moreover, taking into consideration the discretization effects of the flight computer in the design phase improves considerably the stability margins of the controllers. Lastly, it is verified that the proposed controller structures which are designed in the linear domain perform satisfactorily in the nonlinear simulation model and comply with the requirements defined. The multi-modeling controller proved successful in terms of robustness and performance across the flight envelope. Hence, we conclude that implementing this H-infinity optimization process for the design is a promising and viable method to guarantee robustness against uncertainties, disturbances, and measurement noise, as well as compliance with Level 1 HQ. The present work naturally paves the way for recommendations for future work. Some of these recommendations include the extension to lateral-directional control laws, gain-scheduling to guarantee the requirements across the flight envelope, maturity of the aerodynamic data, and pilot assessments of the handling qualities.
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As the aviation industry faces growing environmental and societal pressures, novel concepts such as the Flying-V emerge. Nevertheless, its unconventional configuration poses unique challenges in terms of stability and control, highlighting the pressing need for advanced control systems development. A gap in the literature arises concerning the lack of system robustness in the presence of uncertainties and the absence of robust controllers designed for the aircraft. Hence, the current study aims to increase the maturity of the flight control system of the Flying-V concept by implementing a C* longitudinal controller, while guaranteeing robustness stability and performance against uncertainty, adequate performance in the presence of disturbances and measurement noise, and compliance with Level 1 handling qualities. The Flying-V model is implemented in MATLAB/Simulink which served as the foundation for the controllers synthesis. Three designs are developed, which include a continuous time, a modified continuous time, and a modified continuous time multi-modeling controllers. These are conducted within the standard and signal-based H-infinity mixed sensitivity frameworks. Extensive analysis of the controllers is performed in terms of stability assessments, linear and nonlinear time domain simulations, and uncertainty sensitivity. Results demonstrate that the feedback of a combined signal demand such as the C* parameter provides, for a single input multiple-output system, a balanced disturbance rejection at the plant outputs. Conclusions are drawn in terms of the feedback controller structure, highlighting that high gain is necessary at low frequencies for disturbance rejection and roll-off at high frequencies allows control signal reduction and measurement noise attenuation. Moreover, taking into consideration the discretization effects of the flight computer in the design phase improves considerably the stability margins of the controllers. Lastly, it is verified that the proposed controller structures which are designed in the linear domain perform satisfactorily in the nonlinear simulation model and comply with the requirements defined. The multi-modeling controller proved successful in terms of robustness and performance across the flight envelope. Hence, we conclude that implementing this H-infinity optimization process for the design is a promising and viable method to guarantee robustness against uncertainties, disturbances, and measurement noise, as well as compliance with Level 1 HQ. The present work naturally paves the way for recommendations for future work. Some of these recommendations include the extension to lateral-directional control laws, gain-scheduling to guarantee the requirements across the flight envelope, maturity of the aerodynamic data, and pilot assessments of the handling qualities.
A decades-long steady rise in air movements has disproportionally impacted airport-neighbouring communities. These communities bear the brunt of the aviation industry's air, and noise pollution, detrimental to their livelihood. Airports near densely populated areas face increasing societal and governmental scrutiny to reduce their noise pollution and improve the quality of life for airport-neighbouring communities. Airframe noise has become a prominent noise source during approach due to the increased usage of quieter high-bypass ratio turbofan engines. Flyover measurements suggest that the leading edge junction between the fuselage, wing, and slat is a high-intensity localised noise source. This airframe region has been sparsely researched. Therefore, this study aims to describe the relation between the aero-structural design of the fuselage-wing-slat junction and its aeroacoustic footprint for an aircraft in approach configuration.
The research set-up was designed for open-jet numerical simulations using the commercial Lattice Boltzmann Method (LBM) based CFD solver PowerFLOW, with future compatibility for experimental open-jet wind tunnel validation studies at the TU Delft's Anechoic wind tunnel. Three variations of a two-sideplate research set-up were created based on a slat-and-main-wing modified 30P30N airfoil cross-section. A `No Gap' (NG) geometry that connects the slat to both sideplates; a `No Horn & Step Stump' (NH) geometry that has a simplified slat side-edge and slat stump, a feature that blends the main wing with the fuselage; and a `Horn & Smooth Stump' (H) geometry, which incorporates a slat horn on the slat side-edge and a slat stump modelled after the Airbus A320. For all three variations, the relative positions of the slat track, slat side-edge and other junction surfaces were modelled after the Airbus A320 as well. All model variants were analysed based on their far-field noise radiation and the near-field behaviour of aerodynamic turbulent structures.
The results of the simulated scaled models were compared to literature. The noise radiation of the `H' model was scaled in power and frequency (Strouhal-based), and compared to beamforming integration flyover data of a junction of an Airbus A320. The model shows good spectral resemblance for higher frequencies, with larger discrepancies at lower frequencies. The low-frequency discrepancies were attributed to beamforming limitations stemming from Rayleigh's criterion, as well as a likely over-prediction of lower frequency noise by the scaled `H' set-up. Comparison to full-scale Reynolds number tests on a model Airbus A320 revealed the slope of the scaled noise spectrum resembled the slope of the slat noise of the Airbus A320 at full Reynolds number both at low and high frequencies.
Contrary to earlier research, the set-up without a sideplate-slat gap, `NG', produced excess noise compared to the `Gap'-models, `NH' and `H'. The excess noise is attributed to a larger spanwise extent of the slat narrow-band peak noise mechanism. The narrow-band peaks stem from a slat-cusp-to-slat-trailing-edge flow-acoustic resonance. The introduction of a side-edge (for the `NH' and `H' models) increases the spanwise velocity in the slat cove and creates a slat cove wake that gets accelerated over the slat track. This study showed that the altered slat cove wake shape prevents the canonical slat-cusp-to-slat-trailing-edge flow impingement from occurring on part of the slat, thus limiting the extent of the resonance mechanism. This phenomenon was portrayed by visualising the slat cove wake through total pressure isosurfaces and visualising the narrow-band peak noise source regions through a Ffowcs-Williams and Hawkings (FWH)-based source visualisation technique. A lack of loading on the slat (which is a property of the non-modified 30P30N cross-section as well) is hypothesised as the reason why the introduction of a slat side-edge does not increase the noise radiation for the models with a slat side-edge (`NH', and `H'). ...
The research set-up was designed for open-jet numerical simulations using the commercial Lattice Boltzmann Method (LBM) based CFD solver PowerFLOW, with future compatibility for experimental open-jet wind tunnel validation studies at the TU Delft's Anechoic wind tunnel. Three variations of a two-sideplate research set-up were created based on a slat-and-main-wing modified 30P30N airfoil cross-section. A `No Gap' (NG) geometry that connects the slat to both sideplates; a `No Horn & Step Stump' (NH) geometry that has a simplified slat side-edge and slat stump, a feature that blends the main wing with the fuselage; and a `Horn & Smooth Stump' (H) geometry, which incorporates a slat horn on the slat side-edge and a slat stump modelled after the Airbus A320. For all three variations, the relative positions of the slat track, slat side-edge and other junction surfaces were modelled after the Airbus A320 as well. All model variants were analysed based on their far-field noise radiation and the near-field behaviour of aerodynamic turbulent structures.
The results of the simulated scaled models were compared to literature. The noise radiation of the `H' model was scaled in power and frequency (Strouhal-based), and compared to beamforming integration flyover data of a junction of an Airbus A320. The model shows good spectral resemblance for higher frequencies, with larger discrepancies at lower frequencies. The low-frequency discrepancies were attributed to beamforming limitations stemming from Rayleigh's criterion, as well as a likely over-prediction of lower frequency noise by the scaled `H' set-up. Comparison to full-scale Reynolds number tests on a model Airbus A320 revealed the slope of the scaled noise spectrum resembled the slope of the slat noise of the Airbus A320 at full Reynolds number both at low and high frequencies.
Contrary to earlier research, the set-up without a sideplate-slat gap, `NG', produced excess noise compared to the `Gap'-models, `NH' and `H'. The excess noise is attributed to a larger spanwise extent of the slat narrow-band peak noise mechanism. The narrow-band peaks stem from a slat-cusp-to-slat-trailing-edge flow-acoustic resonance. The introduction of a side-edge (for the `NH' and `H' models) increases the spanwise velocity in the slat cove and creates a slat cove wake that gets accelerated over the slat track. This study showed that the altered slat cove wake shape prevents the canonical slat-cusp-to-slat-trailing-edge flow impingement from occurring on part of the slat, thus limiting the extent of the resonance mechanism. This phenomenon was portrayed by visualising the slat cove wake through total pressure isosurfaces and visualising the narrow-band peak noise source regions through a Ffowcs-Williams and Hawkings (FWH)-based source visualisation technique. A lack of loading on the slat (which is a property of the non-modified 30P30N cross-section as well) is hypothesised as the reason why the introduction of a slat side-edge does not increase the noise radiation for the models with a slat side-edge (`NH', and `H'). ...
A decades-long steady rise in air movements has disproportionally impacted airport-neighbouring communities. These communities bear the brunt of the aviation industry's air, and noise pollution, detrimental to their livelihood. Airports near densely populated areas face increasing societal and governmental scrutiny to reduce their noise pollution and improve the quality of life for airport-neighbouring communities. Airframe noise has become a prominent noise source during approach due to the increased usage of quieter high-bypass ratio turbofan engines. Flyover measurements suggest that the leading edge junction between the fuselage, wing, and slat is a high-intensity localised noise source. This airframe region has been sparsely researched. Therefore, this study aims to describe the relation between the aero-structural design of the fuselage-wing-slat junction and its aeroacoustic footprint for an aircraft in approach configuration.
The research set-up was designed for open-jet numerical simulations using the commercial Lattice Boltzmann Method (LBM) based CFD solver PowerFLOW, with future compatibility for experimental open-jet wind tunnel validation studies at the TU Delft's Anechoic wind tunnel. Three variations of a two-sideplate research set-up were created based on a slat-and-main-wing modified 30P30N airfoil cross-section. A `No Gap' (NG) geometry that connects the slat to both sideplates; a `No Horn & Step Stump' (NH) geometry that has a simplified slat side-edge and slat stump, a feature that blends the main wing with the fuselage; and a `Horn & Smooth Stump' (H) geometry, which incorporates a slat horn on the slat side-edge and a slat stump modelled after the Airbus A320. For all three variations, the relative positions of the slat track, slat side-edge and other junction surfaces were modelled after the Airbus A320 as well. All model variants were analysed based on their far-field noise radiation and the near-field behaviour of aerodynamic turbulent structures.
The results of the simulated scaled models were compared to literature. The noise radiation of the `H' model was scaled in power and frequency (Strouhal-based), and compared to beamforming integration flyover data of a junction of an Airbus A320. The model shows good spectral resemblance for higher frequencies, with larger discrepancies at lower frequencies. The low-frequency discrepancies were attributed to beamforming limitations stemming from Rayleigh's criterion, as well as a likely over-prediction of lower frequency noise by the scaled `H' set-up. Comparison to full-scale Reynolds number tests on a model Airbus A320 revealed the slope of the scaled noise spectrum resembled the slope of the slat noise of the Airbus A320 at full Reynolds number both at low and high frequencies.
Contrary to earlier research, the set-up without a sideplate-slat gap, `NG', produced excess noise compared to the `Gap'-models, `NH' and `H'. The excess noise is attributed to a larger spanwise extent of the slat narrow-band peak noise mechanism. The narrow-band peaks stem from a slat-cusp-to-slat-trailing-edge flow-acoustic resonance. The introduction of a side-edge (for the `NH' and `H' models) increases the spanwise velocity in the slat cove and creates a slat cove wake that gets accelerated over the slat track. This study showed that the altered slat cove wake shape prevents the canonical slat-cusp-to-slat-trailing-edge flow impingement from occurring on part of the slat, thus limiting the extent of the resonance mechanism. This phenomenon was portrayed by visualising the slat cove wake through total pressure isosurfaces and visualising the narrow-band peak noise source regions through a Ffowcs-Williams and Hawkings (FWH)-based source visualisation technique. A lack of loading on the slat (which is a property of the non-modified 30P30N cross-section as well) is hypothesised as the reason why the introduction of a slat side-edge does not increase the noise radiation for the models with a slat side-edge (`NH', and `H').
The research set-up was designed for open-jet numerical simulations using the commercial Lattice Boltzmann Method (LBM) based CFD solver PowerFLOW, with future compatibility for experimental open-jet wind tunnel validation studies at the TU Delft's Anechoic wind tunnel. Three variations of a two-sideplate research set-up were created based on a slat-and-main-wing modified 30P30N airfoil cross-section. A `No Gap' (NG) geometry that connects the slat to both sideplates; a `No Horn & Step Stump' (NH) geometry that has a simplified slat side-edge and slat stump, a feature that blends the main wing with the fuselage; and a `Horn & Smooth Stump' (H) geometry, which incorporates a slat horn on the slat side-edge and a slat stump modelled after the Airbus A320. For all three variations, the relative positions of the slat track, slat side-edge and other junction surfaces were modelled after the Airbus A320 as well. All model variants were analysed based on their far-field noise radiation and the near-field behaviour of aerodynamic turbulent structures.
The results of the simulated scaled models were compared to literature. The noise radiation of the `H' model was scaled in power and frequency (Strouhal-based), and compared to beamforming integration flyover data of a junction of an Airbus A320. The model shows good spectral resemblance for higher frequencies, with larger discrepancies at lower frequencies. The low-frequency discrepancies were attributed to beamforming limitations stemming from Rayleigh's criterion, as well as a likely over-prediction of lower frequency noise by the scaled `H' set-up. Comparison to full-scale Reynolds number tests on a model Airbus A320 revealed the slope of the scaled noise spectrum resembled the slope of the slat noise of the Airbus A320 at full Reynolds number both at low and high frequencies.
Contrary to earlier research, the set-up without a sideplate-slat gap, `NG', produced excess noise compared to the `Gap'-models, `NH' and `H'. The excess noise is attributed to a larger spanwise extent of the slat narrow-band peak noise mechanism. The narrow-band peaks stem from a slat-cusp-to-slat-trailing-edge flow-acoustic resonance. The introduction of a side-edge (for the `NH' and `H' models) increases the spanwise velocity in the slat cove and creates a slat cove wake that gets accelerated over the slat track. This study showed that the altered slat cove wake shape prevents the canonical slat-cusp-to-slat-trailing-edge flow impingement from occurring on part of the slat, thus limiting the extent of the resonance mechanism. This phenomenon was portrayed by visualising the slat cove wake through total pressure isosurfaces and visualising the narrow-band peak noise source regions through a Ffowcs-Williams and Hawkings (FWH)-based source visualisation technique. A lack of loading on the slat (which is a property of the non-modified 30P30N cross-section as well) is hypothesised as the reason why the introduction of a slat side-edge does not increase the noise radiation for the models with a slat side-edge (`NH', and `H').
The concept of Distributed Electric Propulsion (DEP) aircraft holds great promise in achieving the goals outlined in Flightpath 2050. This involves utilizing electricity to power propellers and enhancing overall flight performance through interactions between the wings and propellers. This thesis investigates the stability and control aspects of DEP aircraft compared to conventional aircraft. A tool is proposed to estimate the minimum horizontal tail and aileron size for different configurations, considering propeller effects. Analyzing three configurations (turboprop excluding propeller effects, turboprop including propeller effects, and DEP including propeller effects), propellers are found to destabilize the aircraft but enhance take-off rotation and provide additional damping during roll. The DEP aircraft exhibits a 19% smaller normalized horizontal tail size than the turboprop (with propeller effects), attributed to lower propeller destabilizing moments and a shorter fuselage. Despite a lower roll requirement, the DEP aircraft needs the same normalized aileron size due to larger rolling moment of inertia. A sensitivity analysis suggests a T-Tail is optimal for DEP aircraft stability, and adjusting battery placement and reserve fuel improves overall performance.
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The concept of Distributed Electric Propulsion (DEP) aircraft holds great promise in achieving the goals outlined in Flightpath 2050. This involves utilizing electricity to power propellers and enhancing overall flight performance through interactions between the wings and propellers. This thesis investigates the stability and control aspects of DEP aircraft compared to conventional aircraft. A tool is proposed to estimate the minimum horizontal tail and aileron size for different configurations, considering propeller effects. Analyzing three configurations (turboprop excluding propeller effects, turboprop including propeller effects, and DEP including propeller effects), propellers are found to destabilize the aircraft but enhance take-off rotation and provide additional damping during roll. The DEP aircraft exhibits a 19% smaller normalized horizontal tail size than the turboprop (with propeller effects), attributed to lower propeller destabilizing moments and a shorter fuselage. Despite a lower roll requirement, the DEP aircraft needs the same normalized aileron size due to larger rolling moment of inertia. A sensitivity analysis suggests a T-Tail is optimal for DEP aircraft stability, and adjusting battery placement and reserve fuel improves overall performance.
In nature, birds can intelligently adapt their wing shapes to their environment. This paper aims to replicate this capability by designing an online data-driven aerodynamic performance optimization framework for an unconventional morphing aircraft. Compared to state-of-the-art methods, the proposed framework efficiently finds global optima with reduced computational load when addressing time-varying, nonlinear, and non-convex problems. It also demonstrates enhanced adaptability to unforeseen scenarios. In the event of a sudden actuator fault, the algorithm can automatically detect the fault, adapt the onboard data-driven model, and continue performing optimization and trimming tasks using the remaining healthy actuators. Additionally, the paper addresses the optimal number of actuators within a morphing surface, considering the tradeoff between optimization performance and the weight penalty. High-fidelity simulations demonstrate that through active morphing, the proposed framework achieves drag reductions of 1.9–4.9 % during cruise and up to 12.6 % at higher operational lift coefficients (due to heavier weight and lower speed), resulting in an overall drag reduction of 2.98 % over a typical flight cycle, which corresponds to fuel savings of approximately 150 kg/h. This research represents a significant advancement in sustainable aviation, contributing to reduced fuel consumption, lower emissions, and improved fault tolerance for next-generation aircraft.
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In nature, birds can intelligently adapt their wing shapes to their environment. This paper aims to replicate this capability by designing an online data-driven aerodynamic performance optimization framework for an unconventional morphing aircraft. Compared to state-of-the-art methods, the proposed framework efficiently finds global optima with reduced computational load when addressing time-varying, nonlinear, and non-convex problems. It also demonstrates enhanced adaptability to unforeseen scenarios. In the event of a sudden actuator fault, the algorithm can automatically detect the fault, adapt the onboard data-driven model, and continue performing optimization and trimming tasks using the remaining healthy actuators. Additionally, the paper addresses the optimal number of actuators within a morphing surface, considering the tradeoff between optimization performance and the weight penalty. High-fidelity simulations demonstrate that through active morphing, the proposed framework achieves drag reductions of 1.9–4.9 % during cruise and up to 12.6 % at higher operational lift coefficients (due to heavier weight and lower speed), resulting in an overall drag reduction of 2.98 % over a typical flight cycle, which corresponds to fuel savings of approximately 150 kg/h. This research represents a significant advancement in sustainable aviation, contributing to reduced fuel consumption, lower emissions, and improved fault tolerance for next-generation aircraft.
The approach to sustainable aviation requires effective policy and regulation, informed by accurate environmental research. Aviation emissions inventories form the basis of any environmental assessment, however they are based on simplifications and assumptions that introduce a degree of uncertainty. In terms of robust model development, it is necessary to identify significant sources of uncertainty to determine priorities for improvement.
The results of a global sensitivity analysis identifies 5 key contributors to uncertainty across fuelburn and emissions estimates: BADA Drag and Fuel Flow corrective multipliers, Takeoff Weight, Taxi Thrust and Arrival Inefficiency. Across the emitted species, uncertainty is driven by their respective EI.
EUROCONTROL flight track data allows for an analysis of lateral inefficiencies, and improved distribution fitting. The resulting distributions are more representative of actual flight behavior. The introduction of these improved distributions into the original model result in an overall reduction in fuelburn and emissions estimates.
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The results of a global sensitivity analysis identifies 5 key contributors to uncertainty across fuelburn and emissions estimates: BADA Drag and Fuel Flow corrective multipliers, Takeoff Weight, Taxi Thrust and Arrival Inefficiency. Across the emitted species, uncertainty is driven by their respective EI.
EUROCONTROL flight track data allows for an analysis of lateral inefficiencies, and improved distribution fitting. The resulting distributions are more representative of actual flight behavior. The introduction of these improved distributions into the original model result in an overall reduction in fuelburn and emissions estimates.
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The approach to sustainable aviation requires effective policy and regulation, informed by accurate environmental research. Aviation emissions inventories form the basis of any environmental assessment, however they are based on simplifications and assumptions that introduce a degree of uncertainty. In terms of robust model development, it is necessary to identify significant sources of uncertainty to determine priorities for improvement.
The results of a global sensitivity analysis identifies 5 key contributors to uncertainty across fuelburn and emissions estimates: BADA Drag and Fuel Flow corrective multipliers, Takeoff Weight, Taxi Thrust and Arrival Inefficiency. Across the emitted species, uncertainty is driven by their respective EI.
EUROCONTROL flight track data allows for an analysis of lateral inefficiencies, and improved distribution fitting. The resulting distributions are more representative of actual flight behavior. The introduction of these improved distributions into the original model result in an overall reduction in fuelburn and emissions estimates.
The results of a global sensitivity analysis identifies 5 key contributors to uncertainty across fuelburn and emissions estimates: BADA Drag and Fuel Flow corrective multipliers, Takeoff Weight, Taxi Thrust and Arrival Inefficiency. Across the emitted species, uncertainty is driven by their respective EI.
EUROCONTROL flight track data allows for an analysis of lateral inefficiencies, and improved distribution fitting. The resulting distributions are more representative of actual flight behavior. The introduction of these improved distributions into the original model result in an overall reduction in fuelburn and emissions estimates.
This report covers the investigation of impact of uncertainties on the multidisciplinary design optimization of a medium-range single-aisle turbofan aircraft for minimum global warming impact. The employed workflow for the investigation is a five step process, starting with the implementation of the deterministic climate impact model and carrying out of the design optimization for minimal climate impact. The second step involves the characterisation of uncertainties, where the uncertainties within the climate impact model are identified and quantified. The third step involves the uncertainty analysis, where Monte Carlo simulations are performed to estimate the variability in the average temperature reduction potential of the climate-optimized aircraft with respect to the cost-optimized aircraft. In the fourth step, a robust design optimization is carried out using a non-sorted genetic algorithm to minimize both the average temperature response and variability in average temperature response potential. The sensitivity analysis is carried as the last step using the Morris and variance-based Sobol methods, to identify what the key uncertain parameters are towards the uncertainty in climate impact of the aircraft designs. Scientific uncertainty is identified within the linear climate impact model for the carbon impulse response function parameters, species radiative efficiencies, the NOx and contrail altitude forcing factors, methane lifetime, species efficacies, and are all assigned a probabilistic description. Scenario uncertainty is identified in the future average global CO2 atmospheric concentration projection, for which different realistic future scenarios are characterised. Carrying out the uncertainty analysis has shown that the average temperature response reduction potential of the climate-optimized aircraft is highly uncertain, having a 90% likelihood ranging between 17 and 98 % of the average temperature response of the cost optimized aircraft. This is primarily due to large variability in the estimation of contrail average temperature response. Although the robustness-based optimization did not allow to find any significant improvement in robustness for the climate-optimized aircraft, it did allow to identify an array of robust climate-optimized design solutions. From the sensitivity analysis, it was found that the uncertain parameters showing predominant influence on the output variability are the contrail-related radiaitive efficiency and forcing factors. Additionally, a variability of ±50% in average temperature response apportioned to CO2 emissions was identified due to uncertainty related to future average global CO2 concentration projections.
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This report covers the investigation of impact of uncertainties on the multidisciplinary design optimization of a medium-range single-aisle turbofan aircraft for minimum global warming impact. The employed workflow for the investigation is a five step process, starting with the implementation of the deterministic climate impact model and carrying out of the design optimization for minimal climate impact. The second step involves the characterisation of uncertainties, where the uncertainties within the climate impact model are identified and quantified. The third step involves the uncertainty analysis, where Monte Carlo simulations are performed to estimate the variability in the average temperature reduction potential of the climate-optimized aircraft with respect to the cost-optimized aircraft. In the fourth step, a robust design optimization is carried out using a non-sorted genetic algorithm to minimize both the average temperature response and variability in average temperature response potential. The sensitivity analysis is carried as the last step using the Morris and variance-based Sobol methods, to identify what the key uncertain parameters are towards the uncertainty in climate impact of the aircraft designs. Scientific uncertainty is identified within the linear climate impact model for the carbon impulse response function parameters, species radiative efficiencies, the NOx and contrail altitude forcing factors, methane lifetime, species efficacies, and are all assigned a probabilistic description. Scenario uncertainty is identified in the future average global CO2 atmospheric concentration projection, for which different realistic future scenarios are characterised. Carrying out the uncertainty analysis has shown that the average temperature response reduction potential of the climate-optimized aircraft is highly uncertain, having a 90% likelihood ranging between 17 and 98 % of the average temperature response of the cost optimized aircraft. This is primarily due to large variability in the estimation of contrail average temperature response. Although the robustness-based optimization did not allow to find any significant improvement in robustness for the climate-optimized aircraft, it did allow to identify an array of robust climate-optimized design solutions. From the sensitivity analysis, it was found that the uncertain parameters showing predominant influence on the output variability are the contrail-related radiaitive efficiency and forcing factors. Additionally, a variability of ±50% in average temperature response apportioned to CO2 emissions was identified due to uncertainty related to future average global CO2 concentration projections.
The Flying-V is invented as a sustainable innovation for the aviation industry, promising 20% more fuel efficiency than its benchmark, the Airbus A350. There are many challenges to be addressed to bring the Flying-V to a more developed stage, one of them is the design of the structure that supports both the engine and the landing gear, which motivated this research.
The aim of this research was to develop an automated design methodology for unconventional engine mounting structures that can be applied to the Flying-V.
For the Flying-V, three concepts were evaluated, and the selected concept is a skin-stiffened ortho-grid box structure that combines the landing gear bay and the pylon functions. A total of 25 load cases are derived from the certification specifications CS-25, from which 7 critical load cases are selected and used for design optimisation. Preliminary dynamic loads for the landing gear are estimated, and the landing gear parameters are optimised by minimising the load acting on the landing gear bay for different landing load cases, including lateral and one-gear landing. Dynamic loads from the engine arising from an imbalance due to different failure scenarios related to fan blades being lost; the worst failure scenario was used in the design. The selected concept is assembled into a parameterised finite element model containing 30 design variables. The design is optimised using a design of experiments created using Latin Hypercube Sampling method. Four failure modes are computed using a combination of finite element outputs and (semi-)analytical equations accounting for structural instabilities.
The obtained result is a parameterisation of the primary engine mounting structure of the Flying-V, having a structural mass of 3411 kg per half-wing-fuselage and a minimum margin of safety of 0.09 concerning column buckling of a stiffener. This critical margin of safety is found in the lateral landing load case, which introduces a significant moment on the structure. Moreover, the second most critical failure is produced by the maximum thrust of the engine during take-off.
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The aim of this research was to develop an automated design methodology for unconventional engine mounting structures that can be applied to the Flying-V.
For the Flying-V, three concepts were evaluated, and the selected concept is a skin-stiffened ortho-grid box structure that combines the landing gear bay and the pylon functions. A total of 25 load cases are derived from the certification specifications CS-25, from which 7 critical load cases are selected and used for design optimisation. Preliminary dynamic loads for the landing gear are estimated, and the landing gear parameters are optimised by minimising the load acting on the landing gear bay for different landing load cases, including lateral and one-gear landing. Dynamic loads from the engine arising from an imbalance due to different failure scenarios related to fan blades being lost; the worst failure scenario was used in the design. The selected concept is assembled into a parameterised finite element model containing 30 design variables. The design is optimised using a design of experiments created using Latin Hypercube Sampling method. Four failure modes are computed using a combination of finite element outputs and (semi-)analytical equations accounting for structural instabilities.
The obtained result is a parameterisation of the primary engine mounting structure of the Flying-V, having a structural mass of 3411 kg per half-wing-fuselage and a minimum margin of safety of 0.09 concerning column buckling of a stiffener. This critical margin of safety is found in the lateral landing load case, which introduces a significant moment on the structure. Moreover, the second most critical failure is produced by the maximum thrust of the engine during take-off.
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The Flying-V is invented as a sustainable innovation for the aviation industry, promising 20% more fuel efficiency than its benchmark, the Airbus A350. There are many challenges to be addressed to bring the Flying-V to a more developed stage, one of them is the design of the structure that supports both the engine and the landing gear, which motivated this research.
The aim of this research was to develop an automated design methodology for unconventional engine mounting structures that can be applied to the Flying-V.
For the Flying-V, three concepts were evaluated, and the selected concept is a skin-stiffened ortho-grid box structure that combines the landing gear bay and the pylon functions. A total of 25 load cases are derived from the certification specifications CS-25, from which 7 critical load cases are selected and used for design optimisation. Preliminary dynamic loads for the landing gear are estimated, and the landing gear parameters are optimised by minimising the load acting on the landing gear bay for different landing load cases, including lateral and one-gear landing. Dynamic loads from the engine arising from an imbalance due to different failure scenarios related to fan blades being lost; the worst failure scenario was used in the design. The selected concept is assembled into a parameterised finite element model containing 30 design variables. The design is optimised using a design of experiments created using Latin Hypercube Sampling method. Four failure modes are computed using a combination of finite element outputs and (semi-)analytical equations accounting for structural instabilities.
The obtained result is a parameterisation of the primary engine mounting structure of the Flying-V, having a structural mass of 3411 kg per half-wing-fuselage and a minimum margin of safety of 0.09 concerning column buckling of a stiffener. This critical margin of safety is found in the lateral landing load case, which introduces a significant moment on the structure. Moreover, the second most critical failure is produced by the maximum thrust of the engine during take-off.
The aim of this research was to develop an automated design methodology for unconventional engine mounting structures that can be applied to the Flying-V.
For the Flying-V, three concepts were evaluated, and the selected concept is a skin-stiffened ortho-grid box structure that combines the landing gear bay and the pylon functions. A total of 25 load cases are derived from the certification specifications CS-25, from which 7 critical load cases are selected and used for design optimisation. Preliminary dynamic loads for the landing gear are estimated, and the landing gear parameters are optimised by minimising the load acting on the landing gear bay for different landing load cases, including lateral and one-gear landing. Dynamic loads from the engine arising from an imbalance due to different failure scenarios related to fan blades being lost; the worst failure scenario was used in the design. The selected concept is assembled into a parameterised finite element model containing 30 design variables. The design is optimised using a design of experiments created using Latin Hypercube Sampling method. Four failure modes are computed using a combination of finite element outputs and (semi-)analytical equations accounting for structural instabilities.
The obtained result is a parameterisation of the primary engine mounting structure of the Flying-V, having a structural mass of 3411 kg per half-wing-fuselage and a minimum margin of safety of 0.09 concerning column buckling of a stiffener. This critical margin of safety is found in the lateral landing load case, which introduces a significant moment on the structure. Moreover, the second most critical failure is produced by the maximum thrust of the engine during take-off.
The aviation sector is growing every year with a steady pace of 5%. With this growth, the greenhouse emissions can only rise in the future until new innovative aircraft are designed such as electrical aircraft. An interim solution is to switch the fuel from kerosene to Liquefied Natural Gas (LNG). LNG is much cleaner for the environment, cheaper and more abundant. This study looks into retrofitting conventional aircraft to work on LNG, in this case an Airbus A320 since it the most used airplane in the commercial aviation. A cryogenic tank to store the LNG, was designed and incorporated into the cargo bay replacing some of the cargo. The two aircraft were generated and compared using a preliminary design tool, the Initiator. The LNG aircraft achieved a significant reduction in CO2 and NOx , 24% and 69% respectively. In addition to a potential saving of 17% in direct operating cost.
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The aviation sector is growing every year with a steady pace of 5%. With this growth, the greenhouse emissions can only rise in the future until new innovative aircraft are designed such as electrical aircraft. An interim solution is to switch the fuel from kerosene to Liquefied Natural Gas (LNG). LNG is much cleaner for the environment, cheaper and more abundant. This study looks into retrofitting conventional aircraft to work on LNG, in this case an Airbus A320 since it the most used airplane in the commercial aviation. A cryogenic tank to store the LNG, was designed and incorporated into the cargo bay replacing some of the cargo. The two aircraft were generated and compared using a preliminary design tool, the Initiator. The LNG aircraft achieved a significant reduction in CO2 and NOx , 24% and 69% respectively. In addition to a potential saving of 17% in direct operating cost.
Wing Optimisation for Tractor Propeller Configurations
Validation and Application of Low-Order Numerical Models Adapted to Include Propeller-Induced Velocities
Even though propellers are the oldest form of propulsion, they are still a popular choice for unmanned aerial vehicles (UAVs) and passenger aircraft in certain market segments. If the propellers are mounted on the wing, strong propeller-wing interactions alter the aerodynamic efficiency of the aircraft. Using low-order numerical models, it is shown in literature that the wing chord and twist distribution can be changed to maximise this efficiency. These results are only theoretical. This thesis, therefore, aims to validate and apply a numerical model for optimisation of the wing design taking propeller-wing interactions into account.
A vortex lattice method (VLM) was adapted to include the effects of propeller-induced velocities. Comparing the results of the adapted VLM with existing experimental data already validates the numerical model for predicting the lift distribution. To validate it for changes in lift distribution due to wing design changes, a wind-tunnel experiment is set up. Two wings are tested in a tractor propeller configuration. The only difference between the wings is in the twist distribution. To find the lift distribution, the circulation is evaluated in the flow around a wing at several stations along the wingspan. Particle-image velocimetry was used to obtain this flow field. Indeed, the lift distributions measured on both wings are matched by predictions from the adapted VLM, which proofs the numerical model is suitable for qualitative optimisation studies.
For the wing and operating conditions used for the wind-tunnel experiment, an optimisation study is performed using the adapted VLM. It shows the drag can be reduced with 34% by adopting the optimal chord and twist distribution. Even though the operating conditions are not representative for full-scale aircraft or UAVs, it does show there is a great potential for taking propeller-wing interactions into account for the design of the wing. ...
A vortex lattice method (VLM) was adapted to include the effects of propeller-induced velocities. Comparing the results of the adapted VLM with existing experimental data already validates the numerical model for predicting the lift distribution. To validate it for changes in lift distribution due to wing design changes, a wind-tunnel experiment is set up. Two wings are tested in a tractor propeller configuration. The only difference between the wings is in the twist distribution. To find the lift distribution, the circulation is evaluated in the flow around a wing at several stations along the wingspan. Particle-image velocimetry was used to obtain this flow field. Indeed, the lift distributions measured on both wings are matched by predictions from the adapted VLM, which proofs the numerical model is suitable for qualitative optimisation studies.
For the wing and operating conditions used for the wind-tunnel experiment, an optimisation study is performed using the adapted VLM. It shows the drag can be reduced with 34% by adopting the optimal chord and twist distribution. Even though the operating conditions are not representative for full-scale aircraft or UAVs, it does show there is a great potential for taking propeller-wing interactions into account for the design of the wing. ...
Even though propellers are the oldest form of propulsion, they are still a popular choice for unmanned aerial vehicles (UAVs) and passenger aircraft in certain market segments. If the propellers are mounted on the wing, strong propeller-wing interactions alter the aerodynamic efficiency of the aircraft. Using low-order numerical models, it is shown in literature that the wing chord and twist distribution can be changed to maximise this efficiency. These results are only theoretical. This thesis, therefore, aims to validate and apply a numerical model for optimisation of the wing design taking propeller-wing interactions into account.
A vortex lattice method (VLM) was adapted to include the effects of propeller-induced velocities. Comparing the results of the adapted VLM with existing experimental data already validates the numerical model for predicting the lift distribution. To validate it for changes in lift distribution due to wing design changes, a wind-tunnel experiment is set up. Two wings are tested in a tractor propeller configuration. The only difference between the wings is in the twist distribution. To find the lift distribution, the circulation is evaluated in the flow around a wing at several stations along the wingspan. Particle-image velocimetry was used to obtain this flow field. Indeed, the lift distributions measured on both wings are matched by predictions from the adapted VLM, which proofs the numerical model is suitable for qualitative optimisation studies.
For the wing and operating conditions used for the wind-tunnel experiment, an optimisation study is performed using the adapted VLM. It shows the drag can be reduced with 34% by adopting the optimal chord and twist distribution. Even though the operating conditions are not representative for full-scale aircraft or UAVs, it does show there is a great potential for taking propeller-wing interactions into account for the design of the wing.
A vortex lattice method (VLM) was adapted to include the effects of propeller-induced velocities. Comparing the results of the adapted VLM with existing experimental data already validates the numerical model for predicting the lift distribution. To validate it for changes in lift distribution due to wing design changes, a wind-tunnel experiment is set up. Two wings are tested in a tractor propeller configuration. The only difference between the wings is in the twist distribution. To find the lift distribution, the circulation is evaluated in the flow around a wing at several stations along the wingspan. Particle-image velocimetry was used to obtain this flow field. Indeed, the lift distributions measured on both wings are matched by predictions from the adapted VLM, which proofs the numerical model is suitable for qualitative optimisation studies.
For the wing and operating conditions used for the wind-tunnel experiment, an optimisation study is performed using the adapted VLM. It shows the drag can be reduced with 34% by adopting the optimal chord and twist distribution. Even though the operating conditions are not representative for full-scale aircraft or UAVs, it does show there is a great potential for taking propeller-wing interactions into account for the design of the wing.
Ground based experiments are used to understand and measure rotor and airframe aerodynamic performance. However, these experiments have certain limitations. The effect of these limitations are evaluated using computational fluid dynamic (CFD) modeling techniques. Through this study, the 7- by 10-Foot Wind Tunnel experiments of the Large Civil Tilt Rotor (LCTR) at NASA Ames Research Center will be validated using CFD. The CFD tool, RotCFD, that will be used for this study is developed in corporation with The Aeromechanics Branch at NASA Ames Research Center. RotCFD is a RANS solver wherein the fluid flow is governed by the incompressible, laminar Navier-Stokes equations, and a k-ɛ turbulence model. The current blockage study investigates the effect of the blockage generated by the test hardware on the walls by comparing CFD predictions of the LCTR2 model with and without wind tunnel walls to the wind tunnel test data. Furthermore, attention is paid to the side wall pressure distributions due to a large blockage in the tunnel (particularly at yaw angles approaching 20 deg.). Also, a comparison is made between the pressures measured at the pressure ring locations in the settling chamber upstream of the test section for blockage effects. To investigate this problem, a simplified geometry is analyzed for blockage effects in order to see if these different geometries can represent the LCTR as a simplified case to reduce computational time and obtain a quick understanding of tunnel blockage effects. The focus of this research will be to understand the limitations and accuracy of the recent small-scale Large Civil Tilt Rotor (LCTR) wind tunnel test campaigns.
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Ground based experiments are used to understand and measure rotor and airframe aerodynamic performance. However, these experiments have certain limitations. The effect of these limitations are evaluated using computational fluid dynamic (CFD) modeling techniques. Through this study, the 7- by 10-Foot Wind Tunnel experiments of the Large Civil Tilt Rotor (LCTR) at NASA Ames Research Center will be validated using CFD. The CFD tool, RotCFD, that will be used for this study is developed in corporation with The Aeromechanics Branch at NASA Ames Research Center. RotCFD is a RANS solver wherein the fluid flow is governed by the incompressible, laminar Navier-Stokes equations, and a k-ɛ turbulence model. The current blockage study investigates the effect of the blockage generated by the test hardware on the walls by comparing CFD predictions of the LCTR2 model with and without wind tunnel walls to the wind tunnel test data. Furthermore, attention is paid to the side wall pressure distributions due to a large blockage in the tunnel (particularly at yaw angles approaching 20 deg.). Also, a comparison is made between the pressures measured at the pressure ring locations in the settling chamber upstream of the test section for blockage effects. To investigate this problem, a simplified geometry is analyzed for blockage effects in order to see if these different geometries can represent the LCTR as a simplified case to reduce computational time and obtain a quick understanding of tunnel blockage effects. The focus of this research will be to understand the limitations and accuracy of the recent small-scale Large Civil Tilt Rotor (LCTR) wind tunnel test campaigns.
Multiscale Modelling of Discontinuities
Towards accurate computations of shock-turbulence interactions
Master thesis
(2014)
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Jan van Langenhove, Steven Hulshoff, Marc Gerritsma, Martin Ruess, Roelof Vos