Autonomous aircraft platform with electric and hybrid-electric propulsion

Abstract

A hybrid-electric autonomous aircraft includes a fixed-wing airframe sized to carry at least one standard unit load device, a distributed electric propulsion system including one or more electric propellers disposed along a leading edge of the wing, and a hybrid power system including at least one turbogenerator and one or more battery packs to drive the one or more electric propellers.

Description

Attorney Docket No. RUNE-001 WOAUTONOMOUS AIRCRAFT PLATFORM WITH ELECTRIC AND HYBRIDELECTRIC PROPULSIONCROSS-REFERENCE TO RELATED APPLICATION

[0001] This application claims the benefit of and priority to U.S. Provisional Patent Application No. 63 / 677,251 filed on July 30, 2024, the entire disclosure of which is hereby incorporated by reference in its entirety for all purposes.TECHNICAL FIELD

[0002] The application relates generally to the design and development of aircraft, and more specifically to methods and systems for design, development and integration of hybridelectric propulsion for a fixed-wing aircraft.BACKGROUND OF THE DISCLOSURE

[0003] State-of-the-art processes for designing and delivering aircraft are lengthy, cumbersome, and expensive. In addition, currently designed aircraft are generally limited in energy efficiency, which thus place a detrimental effect on environment. Further, the currently designed aircraft also face limitations in loading capacity due to certain unused parts (such as wings). Further limitations of the currently designed aircraft include, but are not limited to, the complicated vehicle operations, which require well-trained pilots and thus make the running of the aircraft business costly and highly priced.

[0004] The foregoing examples of the related art and limitations therewith are intended to be illustrative and not exclusive, and are not admitted to be “prior art.” Other limitations of the related art will become apparent to those of skill in the art upon a reading of the specification and a study of the drawings.SUMMARY OF THE DISCLOSURE

[0005] The present disclosure addresses the above mentioned problems and other problems in the design and development of aircraft by providing a novel aircraft concept that will realize transformational improvements in energy efficiency and environmental impact, coupled with rapid time-to-market through an initial focus on unmanned aerial vehicles utilized for cargo transportation.

[0006] In one aspect, the present disclosure provides a hybrid-electric autonomous aircraft that includes a fixed-wing airframe sized to carry at least one standard unit load1IPTS / 2OOO81736.1Attorney Docket No. RUNE-001 WO device, a distributed electric propulsion system including one or more electric propellers disposed along a leading edge of the wing, and a hybrid power system including at least one turbogenerator and one or more battery packs to drive the one or more electric propellers.

[0007] In another aspect, the present disclosure provides a method for designing a hybridelectric aircraft, and the method includes determining an optimal configuration of the hybridelectric aircraft based on mission energy balance, stability, and propulsion sizing constraints by using a multi-disciplinary design analysis and optimization (MDAO) framework; implementing robust flight control laws that enable a human-in-the-loop (HITL) or human- on-the-loop (HOTL) by using a dynamics and control simulation and algorithm; verifying the HITL or HOTL through operational simulations to simulate remote operator operations; and demonstrating effectiveness of the configuration and HITL or HOTL flight control system architecture through a prototyped cost-effective subscale aircraft.

[0008] The foregoing is a summary and thus contains, by necessity, simplifications, generalizations and omissions of detail; consequently, the summary is illustrative only and is not limiting in any way. Other aspects, inventive features, and advantages of the systems and / or processes described herein may become apparent in the non-limiting detailed description set forth herein.BRIEF DESCRIPTION OF THE DRAWINGS

[0009] These and other features, aspects, and advantages of the present disclosure may become better understood with regard to the following description, and accompanying drawings, where:

[0010] Figure (FIG.) 1 shows an overall technical development approach for designing and constructing the disclosed aircraft, in accordance with one embodiment.

[0011] FIG. 2 shows an example workflow within an aircraft sizing framework, in accordance with an embodiment.

[0012] FIG. 3 shows a specific architecture of a flight mechanics model, in accordance with an embodiment.

[0013] FIG. 4 shows a schematic overview of an autonomous flight management and navigation architecture, in accordance with an embodiment.

[0014] FIG. 5 shows some parametric studies for performance improvements via optimization, in accordance with an embodiment.2IPTS / 2OOO81736.1Attorney Docket No. RUNE-001 WO

[0015] FIGS. 6A-6C collaboratively show a nose study for system integration, in accordance with an embodiment.

[0016] FIG. 7 shows a mission profile, in accordance with an embodiment.

[0017] FIGS. 8A-8B show constraint diagrams for certain constraint analyses, in accordance with an embodiment.

[0018] FIG. 9A shows an early CATIA model of aircraft showing placement of a unit load device (ULD) container, in accordance with an embodiment.

[0019] FIG. 9B shows the side, front, and top views of the designed craft, in accordance with an embodiment.

[0020] FIG. 10 shows a later iteration of the designed aircraft with a slightly larger fuselage to accommodate multiple (2-3) ULD containers for the hybrid version, in accordance with an embodiment.

[0021] FIG. 11 shows an internal layout of a hybrid aircraft, in accordance with an embodiment.

[0022] FIG. 12 shows some of the sizing capabilities of the hybrid configuration, in accordance with an embodiment.

[0023] FIG. 13 shows margins for destabilizing effects, in accordance with an embodiment.

[0024] FIG. 14 shows evidence of longitudinal stability, in accordance with an embodiment.

[0025] FIGS. 15A-15C show the stability of the disclosed aircraft, in accordance with an embodiment.

[0026] FIG. 16A shows the importance of maximizing the lift-to-drag ratio, in accordance with an embodiment.

[0027] FIG. 16B shows the speed vs. range, in accordance with an embodiment.

[0028] FIG. 16C shows where the designed aircraft lies related to other electric and hybrid aircraft in terms of speed and range, in accordance with an embodiment.

[0029] FIG. 17 shows payload vs. Range for full electric aircraft configurations, in accordance with an embodiment.3IPTS / 2OOO81736.1Attorney Docket No. RUNE-001 WO

[0030] FIG. 18 shows the effect of fuselage body diameter on aircraft drag, in accordance with an embodiment.

[0031] FIG. 19 shows an example computing device for implementing systems and methods described in reference to FIGS. 1-18.DETAILED DESCRIPTION

[0032] To make the objectives, features and advantages of the present disclosure more obvious and understandable, the present disclosure will be further described hereinafter with reference to the accompanying drawings and specific embodiments.

[0033] It should be noted that specific details are set forth in the following description to facilitate understanding the present disclosure. However, the present disclosure may be implemented in many other ways different from those described herein, and those skilled in the art may make similar generalizations without violating the connotation of the present disclosure. Therefore, the present disclosure is not limited by the specific embodiments disclosed below.

[0034] The terms used in the embodiments of the present disclosure are merely for the purpose of describing specific embodiments, and are not intended to limit the present disclosure. The singular forms of “a”, “said” and “the” used in the embodiments of the present disclosure and the appended claims are also intended to include plural forms, unless the context clearly indicates other meanings.

[0035] It should be noted that the example embodiments may be implemented in various forms, and should not be construed as being limited to the embodiments set forth herein. On the contrary, the provision of these embodiments makes the present disclosure more comprehensive and complete, and fully conveys the concept of the example embodiments to those skilled in the art. The same reference numerals in the figures indicate the same or similar structures, and thus their repeated description may be omitted. In addition, the similarities between the embodiments may not be repeated.Motivation and Significance

[0036] Disclosed herein include methods and systems for design, development and optimization of a novel aircraft that may realize transformational improvements in energy efficiency and environmental impact, coupled with rapid time-to-market through an initial focus on unmanned aerial vehicles utilized for cargo transportation, which may sidestep some of the regulatory and operational complexities associated with passenger aircraft, allowing for4IPTS / 2OOO81736.1Attorney Docket No. RUNE-001 WO faster development cycles and earlier market entry. In one specific example, a sustainable or synthetic aviation fuel (SAF)-compatible hybrid-electric propulsion for a fixed-wing craft may be constructed through the methods and systems disclosed herein. The constructed aircraft use hybrid-electric distributed propulsion, which strategically combines a fuelburning turbine engine with electric batteries to replace traditional gas turbine engines in fixed wing aircraft.

[0037] The disclosed aircraft show advantages over other existing fixed-wing crafts. For example, by leveraging a hybrid-electric propulsion (HEP) architecture that combines SAF- compatible turbines with electric batteries, the disclosed aircraft show significant improvements in energy efficiency (specific air range, i.e., the distance an aircraft can fly on a given amount of energy or fuel) and reduced environmental impact. This may be achieved by allowing the fuel-burning turbine engine to be sized to correspond to the lower power requirements (e.g., during the cruise phase of flight) while employing batteries to provide additional boost power during more power-intensive phases (e.g., for takeoff and climb). This configuration thus avoids the significant (or even unacceptable) range and payload penalties associated with all-electric propulsion with the current electric battery. The system therefore delivers a better specific air range, meaning a greater distance flown per unit of energy or fuel.

[0038] In addition, by integrating distributed electric propulsion (DEP) that places multiple electronic motors along the wing, the disclosed aircraft allow the use of higher wing loadings that enable higher cruise efficiency (such as lift-to-drag ratio) and cruise performance, while using the accelerated slipstream (i.e., faster-moving airflow created by electric motors and propellers placed around the wing’s leading edge) over the wing to aid the recovery of high-lift performance for low-speed operations. Such aerodynamic advantages reduce takeoff distance and enhance flight control, especially for short-runway or remote operations.

[0039] Further, the disclosed aircraft has human-in-the-loop (HITL) or human-on-the- loop (HOTL) autonomy that balances high onboard automation with remote human oversight. With significant levels of automation existing onboard the remotely piloted or monitored aircraft, the HITL / HOTL autonomy, through its holistic and integrated focus on the design of control laws, inceptors, and remote displays will significantly reduce remote operator workload, operator training time and cost, and ultimately permit a remote operator to supervise the operation of multiple aircraft, drastically improving scalability.5IPTS / 2OOO81736.1Attorney Docket No. RUNE-001 WO

[0040] The additional advantages of disclosed aircraft include their commercial significance and potential in their sustainability impact. For example, a projected 75% reduction in carbon emissions may be achieved through SAF-compatible hybrid-electric propulsion, compared to conventional jet engines fueled by traditional jet fuels. For context, in 2023, global aviation accounted for approximately 1 gigaton (Gt) of carbon emissions, with regional transportation (covering distances of approximately 1000 miles and under) contributing about 25% of this total. A 75% reduction in regional transportation emissions, compounded over a decade, could lead to a cumulative reduction of 1.9 Gt of carbon emissions - an equivalent environmental benefit to planting approximately 67 billion trees.

[0041] The significant reduction in carbon emissions may be driven by the many different factors of the disclosed aircraft. For example, by integrating electric motors with gas turbines, hybrid-electric systems may optimize power generation and distribution, reducing fuel consumption and CO2 emissions, resulting in improved efficiency. In addition, through capturing kinetic energy during descent and braking, and converting it into electrical energy to power an aircraft’s systems, the as-contracted aircraft may minimize energy waste and reduce emissions, facilitating energy recovery. Further, by integrating renewable energy sources, such as SAF, the disclosed aircraft reduce reliance on jet fuels and lower carbon emissions. Furthermore, the disclosed aircraft employ advanced power management algorithms to minimize fuel consumption and emissions while maintaining optimal performance and range by balancing power demands between electric motors and gas turbines.

[0042] Additional advantages of disclosed aircraft include, but are not limited to, lower operation cost. For example, disclosed aircraft and subsequent derivatives (a hybrid-energybased air vehicle platform) may lead to up to 70% reduction in operating costs, with maintenance and energy costs as the key drivers. Hybrid-electric propulsion systems experience less wear and tear due to substantially fewer moving parts, resulting in reduced maintenance frequency and associated costs. Electric motors are more efficient than traditional gas turbine engines, resulting in reduced fuel consumption and lower fuel expenses. The cost of electricity is generally much lower than that of jet fuel, further contributing to operational cost savings. Lastly, by utilizing SAF and potentially leveraging carbon credits, operators may offset fuel costs and further enhance the economic viability of hybrid-electric propulsion systems.6IPTS / 2OOO81736.1Attorney Docket No. RUNE-001 WO

[0043] Other advantages of the disclosed aircraft include certain benefits related to design features like containerization, and autonomy. For example, 4 times improvement in operator loading / unloading time may be achieved through containerization in the design. In addition, autonomy and low-noise operations may unlock a large number of new regional routes globally using underutilized infrastructure, such as smaller airports. Overall, the disclosed aircraft represents a scalable, sustainable, and economically viable step forward in regional aviation and unmanned logistics.

[0044] It is to be noted that the benefits and advantages described herein are not all- inclusive, and many additional features and advantages may be apparent to one of ordinary skill in the art in view of the figures and the following descriptions.Technical Objectives

[0045] The technical objective of the disclosed aircraft is to maximize the performance potential and secure the HITL / HOTL nature of the concept of operations (CONOPS) for the disclosed aircraft. The specific technical objectives include but are not limited to multidisciplinary design optimization (e.g., performing multi-disciplinary design analysis and optimization (MDAO) using an MDAO framework to identify optimal or Pareto-optimal designs that maximize the performance potential of the configuration), stability and control analysis (e.g., determining the “bare airframe” stability characteristics and the control authority of the design and comparing against associated design requirements or guidelines), flight and propulsion system control law development (e.g., developing an HITL / HOTL- centric flight and propulsion control system (FPCS) to manage the vehicle control effectors and hybrid-electric propulsion system), design and simulate subscale prototype (e.g., designing a subscale prototype for flight testing and performing simulated flights using a dynamics and control simulation and algorithm to assess its feasibility). The specific details for achieving these objectives are further described in details below.

[0046] It should be noted that, while subscale prototyping and simulation play a critical role in de-risking and informing the aircraft design, they are not substitutes for full-scale validation. In the disclosed aircraft development, the final aircraft configuration may undergo a comprehensive testing campaign that includes ground testing, flight testing, and systemlevel evaluations. These tests may be configured to validate the accuracy of simulation models, confirm performance targets, and demonstrate safety and reliability under representative operating conditions. The full-scale test program may also support future certification efforts by generating traceable, verifiable data aligned with regulatory7IPTS / 2OOO81736.1Attorney Docket No. RUNE-001 WO requirements and design assurance standards. This staged approach - progressing from analytical modeling to subscale demonstration to full-scale validation - ensures the integrity, performance, and readiness of the aircraft for commercial deployment.Overall Technical Development Approach

[0047] FIG. 1 shows an overall technical development approach 100 for designing and constructing the disclosed aircraft, in accordance with one embodiment. The technical development approach in the illustrated embodiment mirrors the technical innovation of the aircraft. It involves MDAO 110 using an MDAO framework to exploit the full performance and commercial potential of the hybrid-electric distributed propulsion (HEDP) aircraft architecture. This is followed by a dynamics and control simulation and algorithm 120 to implement robust flight control laws that may enable HITL / HOTL. These may be verified through human-in-the-loop operational simulations 130 to simulate remote operator operations. Next, the effectiveness of the aircraft configuration and HITL / HOTL flight control system architecture may be demonstrated at 140 through the use of rapidly prototyped cost-effective subscale aircraft.

[0048] The MDAO framework disclosed herein is configured to facilitate sizing, optimization, and performance analysis of the disclosed unconventional aircraft designs and propulsion system architectures. The framework employs DEP and a hybrid-electric propulsion system architecture. In this regard, some salient features of the MDAO framework are as follows. First, the underlying energy-based approach seeks mission energy balance by treating fuel-burning and electrified propulsion systems in an equivalent manner. Instead of separating the calculations for thermal and electric power sources, the MDAO framework may consider both as forms of energy that can be measured, consumed, stored, and balanced consistently across mission phases. This equivalency allows for direct trade-off and optimization decisions, such as how much energy should come from batteries versus fuel in different flight segments or phases (e.g., climb vs. cruise), and enables performance comparisons across hybrid, electric, and fuel-powered configurations. The MDAO framework disclosed herein facilitates aircraft sizing with all-electric, hybrid-electric, and turbo-electric propulsion system architectures. It is therefore suited for the hybrid-electric propulsion system trades envisioned for the disclosed aircraft.

[0049] Second, the development of disclosed aircraft requires a multi-objective constrained optimization, in which there are multiple potentially competing objective8IPTS / 2OOO81736.1Attorney Docket No. RUNE-001 WO functions, multiple constraints, and a design space where there are continuous aircraft-level design variables (e.g., wing loading, aspect ratio, etc.) as well as discrete variables describing the propulsion system architecture options. The MDAO framework is configured to undertake such multi -objective constrained optimization of electrified propulsion aircraft. For example, the MDAO framework may search through a broad and diverse design space, evaluate how changes to various parameters interact with each other, and identify possible Pareto-optimal solutions. These are designs where no one objective can be improved without worsening another, providing a balanced set of high-performance candidates. This capability is essential for developing next-generation aircraft where multiple performance, efficiency, and sustainability goals must be met concurrently.

[0050] Third, unlike many other conceptual design tools, the MDAO framework does not consider an aircraft to be a point mass. Instead, rotational equilibria are explicitly considered by trimming an aircraft during both the mission analysis and power sizing for both nominal and off-nominal / post-failure flight conditions. In other words, the MDAO framework explicitly calculates not just linear forces (like lift and drag), but also moments (torques) that affect pitching, yawing, and rolling motion. To account for these, the aircraft may be “trimmed” - a process where the flight control system determines the necessary control surface deflections and thrust levels needed to maintain stable flight at each flight condition. This trimming is performed during both the mission-level analysis (e.g., cruise, climb, descent) and during power sizing, including under off-nominal or failure scenarios, such as the loss of one or more distributed propulsors. This level of detail is a key requirement for DEP to capture the potential impact of off-nominal power requirements on propulsion system sizing. In DEP systems disclosed herein, multiple smaller electric motors may be placed across the wings, and any imbalance in thrust (e.g., due to failure or variation) may significantly affect the aircraft’s stability and performance. Accurately modeling these effects may ensure that the propulsion system is properly sized and that the aircraft remains controllable and safe under a wide range of operating conditions.

[0051] Fourth, despite being a conceptual design tool, the MDAO framework has a full geometric representation of the aircraft with geometry export capability. This is significant, as it can model the actual shapes and spatial relationships like wings, fuselage, propellers, and control surfaces. In addition, it may also support geometry export, which allows sized / optimized aircraft definitions coming out of the MDAO framework to be exported to other high-fidelity analysis tools, such as CAD systems or computational fluid dynamics9IPTS / 2OOO81736.1Attorney Docket No. RUNE-001 WO(CFD) solvers, for further analysis. This capability bridges the gap between early-stage design and detailed engineering, enabling a seamless workflow and reducing the need for redundant model recreation.

[0052] Fifth, the MDAO framework allows easy incorporation of additional analysis modules such as stability & control constraints for power sizing, secondary (non-propul si ve) systems architecture, or alternative methods for existing modules such as battery modeling, aerodynamic modeling, or structural weight estimation. This is advantageous as it may allow to use desired aerodynamic tools / databases and customize the underlying analysis, underscoring the modularity and adaptability of the disclosed MDAO framework.

[0053] FIG. 2 shows an example workflow 200 for an aircraft sizing process within the MDAO framework, in accordance with an embodiment. As illustrated in the left part of the figure, there is a setup phase 210 and a data input phase 220 for the aircraft sizing process.

[0054] During the setup phase 210, a key component is the integrated propulsion model (IPM), which serves as the engine for evaluating how different propulsion configurations perform under various design conditions. For example, the IMP model may be built from a combination of lookup tables (e.g., precomputed datasets for quick reference), embedded scripts (e.g., custom computational routines), reduced order models (e.g., simplified versions of complex systems), or response surface equations (e.g., mathematical approximations derived from detailed simulations or experiments). This modularity allows the model to represent both traditional and advanced propulsion systems with a high degree of adaptability.

[0055] Alongside the IPM, weight estimation models are established during the setup phase. These are tailored to the specific aircraft type being designed. For some parts, especially standard or legacy components, statistical weight equations based on historical data are sufficient. However, for more novel elements or when higher accuracy is needed, physics-based methods are used, which derive weight from structural principles, material properties, and load conditions. Accordingly, the weight equations that are employed are vehicle type-specific and are typically a combination of statistical weight equations for some components and physics-based weight estimates for others.

[0056] Geometry parameterization in the setup phase refers to the process of defining an aircraft’s shape and structure using a set of adjustable design variables, so that its geometry can be systematically generated, modified, and optimized during the design process. In the10IPTS / 2OOO81736.1Attorney Docket No. RUNE-001 WO context of the MDAO framework for the disclosed aircraft, geometry parameterization involves specifying key variables that control the physical dimensions and spatial relationships of each aircraft component, such as wing parameters (such as span, area, aspect ratio, taper ratio, sweep angle, airfoil profile), fuselage parameters (such as length, diameter, cross-sectional shape), tail and control surfaces (such as size, location, dihedral angles, etc.) and propulsion components (such as position and size of propellers or motors in DEP).

[0057] Once these parameters are set, geometry update (or sizing) rules ensure that when one parameter changes (e.g., increasing wing span), the rest of the geometry adjusts accordingly to maintain aerodynamic and structural consistency. For example, increasing the wing span while maintaining the same aspect ratio may automatically adjust the wing area and shape. This parameterized approach enables the MDAO tool to rapidly generate and evaluate thousands of design variations, ensure consistency in structural and aerodynamic modeling, and automatically export geometry to CAD or CFD tools for further analysis. In essence, geometry parameterization turns the aircraft into a flexible, algorithmically controlled shape that can be reshaped and refined based on performance goals and constraints, forming a critical part of the optimization process.

[0058] During the input phase 220 of the aircraft sizing process, key technical and operational parameters are defined to guide the optimization of the aircraft design. Briefly, the technology state-of-the-art (SOTA) captures the current capabilities of core propulsion technologies, such as the specific energy and specific power of batteries, the power-to-mass ratio of electric motors, power electronics, turboshaft engines, etc. These inputs serve as technical bounds and baselines for evaluating design feasibility and performance.

[0059] In parallel, the point performance constraints define the operational scenarios under which the aircraft must be evaluated. These include nominal and off-nominal flight conditions that are evaluated while sizing the propulsion system. These constraints ensure the propulsion system is sized not just for optimal performance, but also for resilience under failure conditions.

[0060] Other essential inputs include the payload (i.e., the weight of cargo or passengers the aircraft must carry) and the mission profile (i.e., a detailed description of flight phases, distances, altitudes, etc.). Together, these shape the overall energy and performance requirements of the vehicle.11IPTS / 2OOO81736.1Attorney Docket No. RUNE-001 WO

[0061] One particularly important constraint is the end-of-mission energy reserve, for example, requiring that the battery must still retain at least 20% charge by the end of the mission. This ensures operational safety, accounting for contingencies like unexpected delays or diversions.

[0062] Furthermore, the sizing process of the input phase 220 additionally includes an initial estimate of the maximum takeoff weight (MTOW). Through iterative simulations, this value is adjusted until the aircraft meets all design objectives and constraints, ensuring it is both efficient and operationally viable. According to some embodiments, the MTOW for the disclosed aircraft may be at least 6,000 pounds.

[0063] As also illustrated in FIG. 2, for the internal structure of the disclosed MDAO framework, it includes a vehicle / system updater 230, a mission analyzer 240, and a mass / energy converter 250 disposed between the vehicle / system analyzer and the mission analyzer. The vehicle / system updater 230 is a comprehensive engine that dynamically adjusts the aircraft configuration during sizing iterations. According to the illustrated embodiment, the vehicle / system updater 230 itself further includes a geometry sizer, a power sizer, a component and empty mass calculator, an energy mass calculator, and a center of gravity ( C.G.) and inertia calculator.

[0064] For the geometry sizer, the MDAO framework uses a representation of the aircraft geometry. Specifically, a user / analyst is able to specify the geometry of a component such as a wing in terms of wing-specific design variables such as span, area, aspect ratio, taper ratio, sweep, airfoil profile, etc., from which the geometry engine generates the wing geometry as a 3-dimensional point list. Similar procedures are followed for the geometry of fuselages, propellers / rotors, etc. Through the geometry sizer, the geometry definition may be automatically updated during each sizing iteration based on the geometry update or sizing rules, ensuring consistency and adaptability as the aircraft scales. Notably, the geometry definition of the vehicle may be exported in other industry standard CAD file formats, which may then be imported into multiple tools, enabling seamless integration with high-fidelity design and simulation tools. In one example application, an MDAO framework-generated geometry may be imported into the industry standard commercial CFD solver for further aerodynamic analysis.

[0065] For the power sizer in the vehicle / system updater 230, it is responsible for determining how much power each propulsion system component (such as electric motors,12IPTS / 2OOO81736.1Attorney Docket No. RUNE-001 WO batteries, or turbines) needs to deliver across all relevant flight scenarios. These scenarios include user-defined nominal flight conditions (such as cruising or climbing during routine operations) and off-nominal flight conditions (such as emergency descent or a partial power system failure). At each such condition, the aircraft is trimmed (i.e., a process that adjusts control surfaces and thrust levels to ensure stable flight) and the propulsion system is evaluated through power flow analysis (i.e., a process that tracks how energy is distributed through the propulsion system during that flight condition). By running this process for all defined flight scenarios, the system may identify the most demanding case for each component, the one that requires the most power or performance capacity. This value is known as the constraining or limiting power, and it is what determines the minimum required power size of that component to ensure safe and effective operation under all expected conditions. In essence, the power sizer ensures that every propulsion component is robust enough to handle the worst-case scenario it might face, without being oversized to the point of inefficiency.

[0066] For the component and empty mass calculator, this tool may estimate the mass of individual aircraft components, such as the fuselage, wings, landing gear, propulsion units, avionics, and structural supports. It also sums these to determine the empty mass of the aircraft, which is the weight of the vehicle without payload, fuel, or batteries. The calculations may use a mix of statistical data (e.g., based on historical aircraft) and physicsbased models tailored to the specific configuration and materials of the aircraft. Accurate empty mass estimation is essential for performance analysis and sizing of the propulsion and structural systems.

[0067] For the energy mass calculator, this tool may determine how much mass must be allocated to energy storage, i.e., batteries, fuel tanks, and related systems, to meet the mission energy demands. It translates energy requirements (in kilowatt-hours or joules) into mass, considering the specific energy of each source (e.g., Wh / kg for batteries, MJ / kg for fuel). This step is crucial in hybrid-electric aircraft, where balancing energy capacity with overall vehicle weight significantly impacts range, efficiency, and feasibility.

[0068] For the C.G. and inertia calculator, this tool may compute the aircraft’s center of gravity and moments of inertia based on the locations and masses of all components. C.G. is vital for ensuring longitudinal and lateral stability, while inertia values affect how the aircraft responds to control inputs and external forces. These calculations support stability analysis,13IPTS / 2OOO81736.1Attorney Docket No. RUNE-001 WO flight dynamics modeling, and control system design, and are essential for trimming the aircraft during simulation and sizing.

[0069] Together, these calculators provide a comprehensive physical characterization of the aircraft, enabling precise modeling, optimization, and performance prediction across the entire design process.

[0070] For the mission analyzer 240, it is another module within the MDAO framework responsible for simulating and evaluating how the aircraft performs across its entire mission profile, from takeoff through cruise to landing. It plays a central role in determining whether a given aircraft configuration can successfully complete its assigned mission within the constraints of energy, weight, payload, and performance. According to some embodiments, the mission analyzer may simulate the full mission timeline. For example, the mission analyzer may model each flight segment (e.g., takeoff, climb, cruise, descent, landing) based on the mission profile input by the user. This includes altitude, speed, distance, and duration of each phase. In addition, the mission analyzer may also calculate how much energy is consumed in each segment, drawing from electric and / or fuel-based power sources, thereby ensuring that the total energy required fits within the capacity provided by the energy systems. This includes verifying that the final battery state of charge or fuel level meets end- of-mission constraints. In some embodiments, the mission analyzer may further check performance against constraints. This ensures that key performance metrics, like range, endurance, speed, and climb rate, are met at each stage of the mission, both under nominal and off-nominal conditions (e.g., degraded propulsion performance). In some embodiments, the mission analyzer may provide feedback for optimization. For example, the results from the mission analyzer may be fed back into the MDAO loop, guiding design updates. If the aircraft cannot complete the mission (e.g., run out of energy or exceeds weight limits), parameters such as battery size, fuel capacity, or aerodynamic features are adjusted. In some embodiments, the mission analyzer may support additional trade studies. For example, by running many mission scenarios across different aircraft configurations, the mission analyzer may enable comparative analysis, identifying which design choices yield the best performance, efficiency, or cost-benefit under realistic operational conditions. In short, the mission analyzer ensures that the aircraft not only looks good on paper but can actually fly its intended mission reliably and efficiently, validating the design against real-world demands.

[0071] In some embodiments, the mission analyzer 240 may employ a flight mechanics model (FMM) to run the above described analysis. This model ensures that the aircraft’s14IPTS / 2OOO81736.1Attorney Docket No. RUNE-001 WO aerodynamic performance, energy consumption, and control stability are accurately represented across its mission. The specific detail for the flight mechanics model is further described in FIG. 3.

[0072] FIG. 3 illustrates a comprehensive flight mechanics model 300 developed to simulate and evaluate the performance characteristics of the disclosed electric and hybridelectric cargo aircraft, in accordance with an embodiment. The model serves as a critical analytical framework to assess the dynamic behavior of the aircraft by integrating aerodynamic, geometric, and propulsion-related data. This is especially relevant for the disclosed system, which involves a distributed electric propulsion system and a novel high- aspect-ratio wing configuration designed to maximize lift and minimize drag under varying flight conditions.

[0073] At the front end of the model, aircraft-specific data including aircraft geometry 310 and information from an aero database 320 are supplied. These inputs define the physical shape, aerodynamic surfaces, and performance baselines of the aircraft. This geometric and aerodynamic foundation feeds into the aero-propulsive model 330, which plays a pivotal role in generating key aerodynamic coefficients such as the lift coefficient (CL), drag coefficient (CD), and moment coefficient (cmA). These coefficients are dynamically influenced by the distributed propulsion system’s operation and the high-lift features integrated into the aircraft, including propeller-blown lift and trailing-edge flap systems.

[0074] The derived aerodynamic parameters are subsequently processed through the equations of motions 340, which simulate how the aircraft responds to external forces and internal control inputs. This includes translation and rotation in three dimensions, and factors in aircraft mass properties, thrust-to-weight ratios, and real-time power availability from the battery and / or hybrid systems. The model accounts for various mission segments, such as takeoff, climb, cruise, and landing, and is particularly adept at modeling the complex interactions between distributed propulsion and wing aerodynamics.

[0075] Thereafter, the simulation outputs 350 provide essential insights into aircraft behavior, including stability margins, maneuverability, energy consumption, and field performance (e.g., takeoff roll distances). These outputs guide design refinements and validate that the aircraft meets FAA Part 23 airworthiness standards, even with novel configurations.15IPTS / 2OOO81736.1Attorney Docket No. RUNE-001 WO

[0076] The flight mechanics model depicted in FIG. 3 is thus essential to verifying that the disclosed unique aerodynamic and propulsion systems function cohesively to deliver optimized performance with minimal compromise on stability or range.

[0077] Referring back to FIG. 2, the mass / energy converter 250 may ensure that the aircraft’s energy system is properly balanced for its mission. Specifically, the converter may adjust the MTOW of the aircraft iteratively until it achieves mission energy balance. Mission energy balance means that the total energy available (from both fuel and battery sources) is exactly sufficient to meet the total energy required throughout the mission. This required energy includes not only what’s consumed during flight phases like takeoff, climb, cruise, and descent, but also any mandated energy reserves, such as a minimum state-of-charge for the battery at the end of the mission (e.g., 20%). If the initial MTOW guess is too low, the aircraft may not have enough energy capacity (fuel or battery) to complete the mission. Conversely, if the MTOW is too high, it may exceed performance or efficiency constraints. The mass / energy converter continually adjusts MTOW, alongside related parameters like fuel and battery sizing, until the energy budget is balanced without violating any design constraints. This ensures the aircraft is both operationally viable and energy-efficient for its intended mission profile.

[0078] It is to be noted that at the core of the MDAO framework are the IPM and FPM. The generalized formulation of the IPM is responsible for the demonstrated general applicability of the MDAO framework to various aircraft configurations and propulsion architectures. The FPM computes the trim solution of the aircraft by solving the IPM in one of several modes (power, flightpath, or acceleration).

[0079] Reference is now made to the dynamics and control simulation and algorithm 120, another important framework used in the technical development approach of the disclosed aircraft. This MATLAB / Simulink-based framework is configured to enable the assessment of the stability and control characteristics and the design of the flight control system. Salient features of the dynamics and control simulation and algorithm in this regard are as follows.

[0080] The dynamics and control simulation and algorithm 120 employs a generalized formulation of the trim problem, where both conventional control surfaces (e.g., ailerons, rudder, elevator) and propulsors (e.g., electric motors in distributed electric propulsion) may act as control effectors. This is essential for hybrid-electric or DEP aircraft, where propulsion plays an active role in maneuvering and stability.16IPTS / 2OOO81736.1Attorney Docket No. RUNE-001 WO

[0081] Trim is calculated not through simple force balance, but by solving a constrained minimization problem, where a user-defined objective function is minimized (e.g., required control effort or fuel use) while ensuring all physical and operational constraints are satisfied. The dynamics and control simulation and algorithm allows the trimmability of a disclosed aircraft to be assessed over its flight envelope for both nominal and off-nominal flight conditions.

[0082] After determining a valid trim condition, the dynamics and control simulation and algorithm 102 may numerically linearize the aircraft dynamics model and determine dynamic stability characteristics by solving eigenvalue problems. This analysis may be done both for the “bare airframe” (open-loop) as well as for the closed-loop system with flight control laws in effect. This capability is an enabler for the design of the HITL / HOTL flight control system architecture envisioned for the disclosed aircraft.

[0083] In some embodiments, the disclosed flight control system architecture may be implemented in the dynamics and control simulation and algorithm using a Simulink model, which may allow flight control laws (e.g., the key elements of the HITL / HOTL system) to be tested through time-domain simulations. These simulations, which may be either autonomous or piloted, may be run in real-time or faster than real-time. This enables robust development and testing of autonomous and semi-autonomous control strategies, validating how the aircraft would behave under both normal and degraded conditions, and supporting the architecture of a scalable, safe, and certifiable HITL / HOTL system.Specific Implementations

[0084] The specific implementations of design, development and optimization of the disclosed aircraft may be categorized into specific tasks, such as tasks for aircraft baseline model development in the MDAO framework, multi-disciplinary design analysis and optimization, stability and control assessment in the dynamics and control simulation and algorithm, flight and propulsion control law development in the dynamics and control simulation and algorithm, subscale prototype aircraft design using the MDAO framework, and subscale prototype aircraft simulation using the dynamics and control simulation and algorithm.

[0085] Baseline Model Development

[0086] Specifically, for the task of aircraft baseline model development in the MDAO framework, the starting point may be an existing design iteration of an aircraft that has been17IPTS / 2OOO81736.1Attorney Docket No. RUNE-001 WO obtained through prior analyses or optimization. In addition to the baseline geometry, the task may further include the selection of weight estimation relationships for major structural components, as well as other “known” or off-the-shelf components that will be present in the aircraft, such as standardized avionics, landing gear, or propulsion modules, which may already have known weights, dimensions, and performance characteristics.

[0087] Additional implementations include the incorporation of the aerodynamic database for the aircraft into the strip theory -based IPM within the MDAO framework, enabling accurate performance modeling during flight simulations. It is anticipated that a mix of aerodynamic data generated by mid-to-high fidelity flow solvers and CFD tools may be used within the IPM, ensuring a realistic and reliable representation of aerodynamic behavior under various conditions. In one example, a turbogenerator (hybrid-electric) engine that will power its first full-scale aircraft may be selected. This shows that the baseline model isn't static, it evolves with the incorporation of real or emerging technologies. The selection of propulsion architecture and available technologies also helps narrow and specify the mission definition and performance requirements, thereby guiding subsequent rounds of optimization and simulation. This process forms the bridge between early conceptual design and more detailed, technically grounded engineering development.

[0088] In some embodiments, proprietary engine performance data tables (“engine decks”) for this turbogenerator engine may be incorporated into the MDAO framework to model the power lapse (i.e., how engine power decreases with altitude or temperature) and fuel consumption rates of the engine. Incorporating such data allows the simulation to model the engine’s behavior with high precision throughout the mission profile.

[0089] In addition to performance data, manufacturer data for the uninstalled (engine alone) and installed (with mounts, nacelles, and integration hardware) weight of the engine may be also available. This ensures accurate modeling of the propulsion system’s mass contribution to overall aircraft weight and balance, which is critical for performance prediction, stability, and control analysis. Similarly, a range of battery technology levels defined by specific energy (e.g., energy per unit mass) and specific power (e.g., power per mass unit) are also considered in the MDAO framework. These values are selected based on specifications from available commercial battery vendors. This approach allows designers to explore trade-offs between competing battery technologies, balancing energy capacity, weight, discharge rates, and integration feasibility, within the bounds of real, manufacturable systems.18IPTS / 2OOO81736.1Attorney Docket No. RUNE-001 WO

[0090] According to some embodiments, the completion of the task of aircraft baseline model development in the MDAO framework may yield a baseline model that is modeled in the MDAO framework, which may be fed into the following task (e.g., MDAO) for further processing.

[0091] Multi-disciplinary Design Analysis and Optimization

[0092] For the task of multi-disciplinary design analysis and optimization, following the baseline model development in the MDAO framework, the analysis and optimization may be performed on the design. For example, the non-dominated sorting algorithm II (NSGA-II) genetic algorithm (GA) may be used. There are several advantages of a GA in this context: (i) the GA, by its very design, is naturally able to handle a combination of continuous and discrete design variables - this will allow the HEDP design space to be optimized; (ii) the GA may account for multiple (potentially competing) objective functions without requiring the arbitrary specification of objective weighting functions; (iii) the GA is more robust to the presence of local minima (i.e., solutions that are better than nearby alternatives but not globally optimal) in the design space due to its population-based approach and stochastic nature, which allows to it be better at exploring the entire design space and avoiding this trap; (iv) rather than yielding a single supposedly optimal design, the GA may yield an entire population of Pareto-optimal solutions as the outcome of a multi-objective optimization with two or more competing constraints. This gives engineers flexibility and insight when selecting final configurations, as they can choose among trade-offs that best suit the mission, budget, or technological readiness.

[0093] For the MDAO disclosed herein, continuous variables may include but are not limited to traditional aircraft design variables such as wing loading, aspect ratio, taper ratio, airfoil thickness-to-chord ratio and other airfoil characterization parameters, stabilizer volume ratios, fuselage shape and fineness ratio, etc. Discrete design variables govern the number of distributed propulsors per wing, the number of blades per propeller, the empennage configuration, the number of battery packs, component connectivity within the HEDP architecture, HEDP power management strategies over the course of the mission, etc. Objective functions (lower-is-better type) include but are not limited to maximum takeoff weight (MTOW), empty weight or empty weight fraction, propulsion + energy weight or weight fraction, etc.19IPTS / 2OOO81736.1Attorney Docket No. RUNE-001 WO

[0094] The MDAO disclosed herein may mitigate the technical risk. For example, the MDAO may be run on cluster computers after the parallelization of instances of the MDAO framework, significantly accelerating the optimization process by allowing multiple simulations or design evaluations to run simultaneously. This scalability is essential when exploring large, complex design spaces with many variables and constraints. Additionally, the MDAO may be configured to run across different technology scenarios, for example, varying the specific energy of batteries, the efficiency of electric motors, or the performance of different turboshaft engines. By evaluating these variations systematically, the framework may assess how sensitive the Pareto-optimal solutions are to shifts in technology performance. This kind of sensitivity analysis helps identify which technologies are most critical to achieving performance goals and which areas might introduce the greatest uncertainty or design risk.

[0095] According to some embodiments, the task of MDAO may further include a tradespace analysis, which involves evaluating a wide variety of design options to understand trade-offs between performance, efficiency, and emissions. In the context of aircraft design disclosed herein, a tradespace analysis may use high fidelity CFD combined with blade element momentum (BEM) theory to further complement the above task and optimize the distributed electric propulsion system for the selected wing geometry. The optimization of distributed electric propulsion system is combined with hybrid-electric powertrain optimization to achieve a high lift-to-drag (L / D) ratio during the cruise phase, hence further optimizing fuel burn and emission. In the end, the objective is to reduce fuel bum and emission by a significant amount (e.g., 40% or higher) using the hybrid-electric and DEP system and define a path towards future zero-emission regional travel configuration. The analysis may investigate and optimize multiple mission parameters, such as endurance (i.e., how long the aircraft can fly), payload capability (i.e., how much cargo or passengers it can carry), electric power output (e.g., sizing and performance of onboard energy systems), operating ceiling (i.e., maximum altitude), and maximum speed, as well as multiple alpha and beta angles. Using Reynolds-averaged Navier-Stokes (RANS) solver and an established cloud-based GPU-enabled CFD solution may allow for rapid iterations, optimized configuration and higher performance (payload / range). RANS-based CFD may simulate complex airflow over the entire aircraft, enabling highly accurate predictions of lift, drag, and propulsion efficiency.20IPTS / 2OOO81736.1Attorney Docket No. RUNE-001 WO

[0096] According to some embodiments, the completion of the task of multi-disciplinary design analysis and optimization may yield Pareto-optimal designs for the aircraft, from which one or more viable candidates will be down-selected for further analysis in later tasks.

[0097] Stability and Control Assessment

[0098] With respect to the task of stability and control assessment, the objective is to assess the stability and control (S&C) characteristics of the designs down-selected from the task of multi-disciplinary design analysis and optimization in a greater level of detail. Even though an aircraft is intended to be fly-by-wire (which refers to a modern aircraft control system where pilot inputs are transmitted electronically rather than through traditional mechanical linkages like cables, pulleys, or rods), acceptable or at least adequate “bare airframe” stability characteristics (meaning their natural flight stability without assistance from active control systems like fly-by-wire) are nevertheless desirable. It simplifies control law development, improves safety margins, and enhances robustness. This may be assessed through the dynamic stability of the sized designs using the dynamics and control simulation and algorithm, which analyzes how the aircraft responds to disturbances such as wind gusts or control inputs, and whether it returns to steady flight without diverging. If serious deficiencies are found (e.g., poor pitch damping or unstable roll / yaw behavior), the deficiencies may be reflected back into the task of multi-disciplinary design analysis and optimization, for example, in the form of more stringent constraints on stabilizer sizing (e.g., larger or differently shaped stabilizers) that ensure future iterations of the design have improved stability.

[0099] In some embodiments, in addition to overall dynamic stability, controllability assessment may be also performed, which may include checking for adequate control authority for takeoff rotation (lifting the nose off the ground), landing flare, and crosswind operations at the low speeds anticipated for this blown lift concept. Additional checks may include controllability following the failure of one or more distributed propulsors, ensuring the aircraft remains controllable after the failure of one or more distributed propulsors, and a realistic scenario in DEP systems where many smaller motors are used. The linear time invariant (LTI) state-space models identified during this task may be fed into the following task for the development of the flight and propulsion control systems.

[0100] According to some embodiments, the completion of the task of stability and control assessment in the dynamics and control simulation and algorithm may generate21IPTS / 2OOO81736.1Attorney Docket No. RUNE-001 WO detailed knowledge regarding the stability and control characteristics of the aircraft and generate LTI state-space models that may be used in the task of flight & propulsion control law development in the dynamics and control simulation and algorithm, as described in detail below.

[0101] Autonomous Flight Management and Navigation Architecture

[0102] FIG. 4 illustrates a schematic overview of the autonomous flight management and navigation architecture used in the electric and hybrid-electric aircraft system, in accordance with an embodiment. The figure illustrates the end-to-end decision-making and control flow 400 that enables fully autonomous operation, from mission planning to real-time navigation. This framework is particularly relevant for the described unmanned aerial vehicle (UAV), which does not rely on continuous pilot input but rather on a blend of pre-programmed commands and adaptive onboard intelligence.

[0103] The process 400 begins with a ground operator who initiates the mission plan and / or command at step 410. This operator may oversee multiple aircraft simultaneously in an m:N operational model, meaning one operator can control or monitor multiple aircraft in parallel. The mission plan and / or command is then transmitted via a C2 / C3 (command and control / communications and coordination) link to the aircraft’s guidance, navigation, and control (GNC) system. This subsystem is responsible for converting mission objectives into navigable routes and generating an accurate environmental map of the aircraft’s operational zone through mission to navigation routing and mapping at step 420.

[0104] Once routing and mapping are established, the information flows into the flight control and autonomous control recognition and mapping module at step 430. Here, the aircraft dynamically manages in-flight decision-making, including obstacle avoidance, energy consumption optimization, and coordination with its propulsion and aerodynamic systems. This part of the system ensures that the aircraft can respond to real-world changes and uncertainties, such as weather fluctuations or unexpected obstructions, without requiring immediate operator input.

[0105] Thereafter, the vehicle executes waypoint and / or directional navigation at step 440, adjusting its path based on both pre-defined and dynamically generated waypoints. This phase supports integration with air traffic control (ATC) environments, such that the aircraft can interact with one or more ATC units (ATC 1, 2, ..., n) as needed. It also supports m:N22IPTS / 2OOO81736.1Attorney Docket No. RUNE-001 WO autonomous operation, reinforcing the system’s scalability and suitability for fleet-based logistics or commercial cargo delivery missions.

[0106] FIG. 4 thus encapsulates the intelligent, distributed decision-making system that enables the aircraft to conduct long-range, unmanned missions with high levels of autonomy, situational awareness, and airspace compliance, key aspects of the disclosed aircraft’s operational capabilities.

[0107] Subscale Prototype Aircraft Design and Simulation

[0108] For the task of subscale prototype aircraft design using the MDAO framework, it serves as a crucial intermediate step between simulation and full-scale development. The goal is to create a physically testable prototype that reflects the optimized configuration of the full- scale aircraft and enables validation of key performance assumptions through flight testing. The starting point of this task is a technical recommendation, derived from prior trade studies and MDAO optimization results. These recommendations help determine a most promising architecture to pursue in physical form. The subscale prototype (or subscale demonstrator) is not just a scaled-down version of the aircraft, it is intentionally designed to test the feasibility and behavior of the system in real-world conditions, particularly those simulated earlier in the development process.

[0109] According to some embodiments, the sizing / optimization of subscale aircraft using the MDAO framework is similar to that of a full-scale aircraft. The aircraft is sized based on a set of point performance requirements (such as minimum climb rate, speed, or maneuverability) and a flight endurance (i.e., how long the aircraft must be able to fly on a given energy source) requirement. This ensures the prototype is not only flyable but can meaningfully validate critical performance metrics.

[0110] For the electric subscale demonstrator, the conclusion of this process may yield commercial off-the-shelf electric motors, such as electric motors and batteries that may allow the performance and endurance constraints to be met based on the energy and power constraints. Similarly, if a hybrid-electric configuration is pursued, a hybrid-electric subscale demonstrator may also be sized using the MDAO framework to balance turbine and electric power contributions.

[0111] According to some embodiments, the completion of the task of subscale prototype aircraft design using the MDAO framework may yield designs for fully defined subscale demonstrators that may be constructed and flight-tested.23IPTS / 2OOO81736.1Attorney Docket No. RUNE-001 WO

[0112] For the task of subscale prototype aircraft simulation using the dynamics and control simulation and algorithm, the development of a flight simulation model for a subscale vehicle in the dynamics and control simulation and algorithm is similar to that for a full-scale vehicle. The aircraft geometry, aero-propulsive characteristics, and mass properties will be obtained from the outcome of the task of subscale prototype aircraft design using the MDAO framework. The same FCS architecture used for the full-scale vehicle in the task of flight & propulsion control law development in the dynamics and control simulation and algorithm may be used for the subscale vehicle. The dynamics and control simulation and algorithm may be used to optimize the gain schedule for the subscale vehicle. This gain schedule may be the starting point for subscale flight tests. They will be refined as required based on the vehicle’s handling characteristics during flight testing.

[0113] According to some embodiments, the completion of the task subscale prototype aircraft simulation using the dynamics and control simulation and algorithm may yield a certifiable full-scale model (i.e., complete with geometry, propulsion architecture, and energy systems) that may be constructed and flight-tested. These simulation models may serve as a proof-of-concept and a critical validation tool, allowing to understand the flight characteristics of the vehicles, while reducing risk for full-scale development and helping to refine simulation models with empirical flight data for certification accordingly.Some Example Embodiments

[0114] Iterative trade studies and constraint analysis that shaped the foundational performance parameters of an aircraft design were conducted throughout the aircraft design process disclosed herein. The goal was to establish the optimal combination of aerodynamic and propulsion characteristics, specifically wing loading (W / S) and thrust-to-weight ratio (T / W), that would satisfy mission and performance goals while also accounting for operational flexibility and future scalability.

[0115] Specifically, the wing loading and thrust-to-weight ratio were determined based on the constraint diagrams’ illustration of acceptable combinations that would meet the design goals. Here, constraint diagrams refer to visual tools used in aircraft performance engineering to show the feasible combinations of W / S and T / W that meet various flight requirements. Using these constraint diagrams, the aircraft lift coefficient needed to achieve the climb and takeoff requirements during critical phases like takeoff and climb was24IPTS / 2OOO81736.1Attorney Docket No. RUNE-001 WO determined after adjusting the aerodynamic parameters such as the aspect ratio, wing area, and MTOW.

[0116] A speed vs. range analysis revealed that the aircraft did not need to meet highspeed targets, since it was intended for short-range missions. This allowed the designers to prioritize energy efficiency and payload over top-end speed. Later in the process, a range vs. payload trade study was conducted and it was determined that the payload goal of greater than 20% above the incumbent players (i.e., existing competitors) would be set as a stretch goal for the range. Finally, recognizing that changes in fuselage size could affect drag, a drag vs. cross-sectional area was investigated to better understand the aerodynamic consequences of potential fuselage growth, especially to accommodate larger payloads or system components.

[0117] To formalize these performance requirements, a comprehensive constraint analysis was conducted to determine the optimum T / W ratio and W / S for given flight conditions and requirements. Specifically, five flight kye flight scenarios were implemented: cruise, ground roll for take-off, rate of climb, level turn, and service ceiling. Gudmundsson’s equations, a standard reference in aircraft design, were used to calculate T / W for a given wing loading at each constraint. The required power was then calculated by converting the thrust requirement into brake horsepower (BHP), giving a realistic estimate of the propulsion power needed. Since not all input data were known early on, initial estimations were made for values not known at the beginning stage of the design process, and the constraint analysis was updated to reflect changes during the design process. Requirements set by the initial design parameters were also implemented.

[0118] FIG. 5 shows some parametric studies for performance improvements via optimization, in accordance with an embodiment. Studies show that DEP improved lift by up to 40% during take-off thus reducing take-off runway length. The addition of winglets had the capability to reduce the induced drag by up to 20%, depending on wing design.Additional performance, i.e., lift-to-drag ratio could be obtained via combining CFD and Adjoint methods to design new airfoils specific to the designed aircraft’s mission. It is to be noted that this path is not limited to airfoil sections but can be extended to an entire new wing design that meets constraint requirements. This dynamic, feedback-driven approach allowed the team to adjust the constraint analysis to reflect improvements in component models, technology assumptions, and mission requirements, ensuring a robust and optimized design path.25IPTS / 2OOO81736.1Attorney Docket No. RUNE-001 WO

[0119] A forward fuselage (including the nose) study was further conducted to support system integration and internal volume expansion without sacrificing aerodynamic performance, as shown in FIGS. 6A-6C. The study used CFD simulations with rapid iteration times to evaluate how different nose geometries impacted the aircraft’s aerodynamic behavior, particularly its L / D ratio.

[0120] FIG. 6A shows that CFD simulations with a fast turn-around time enabled quick assessment of various nose designs and impact on performance. Specifically, different nose configurations were tested, with the aim of striking an optimal balance between performance, functionality and aesthetics. FIG. 6A illustrates a comparison of one design concept (i.e., nose concept 2) against a baseline. In the figure, the baseline refers to the conventional rounded nose and is regarded as a conservative design, and concept 2 refers to a flattened nose with a wide cross-sectional width and sharp edge, aimed at increasing internal volume while preserving aerodynamic efficiency. The goal was to strike a practical balance between performance, functionality, and aesthetics.

[0121] FIG. 6B shows the lift-to-drag ratio vs. angle of attack for nose studies. From the figure, it can be seen that, compared to the baseline, none of the alternative designs significantly degraded aerodynamic performance. Even as the nose designs deviated from the baseline, CFD analysis confirmed minimal impact on the aircraft’s overall L / D ratio. This suggests the aircraft could adopt a more space-efficient nose design without compromising efficiency or range.

[0122] FIG. 6C further shows fuselage lift and drag analysis. In the figure, the plot shows a comparative drag and lift ratio of nose concept 2 compared to the reference baseline. Because of the increased surface area for the nose concept 2, the drag increases as expected. In addition, for nose concept 2, the fuselage lift increases more compared to the fuselage drag, thus mitigating the drop in the overall lift-to-drag ratio. In fact, the fuselage for nose concept 2 acts as an additional lifting surface.

[0123] In terms of internal layout and integration, this redesigned nose offered substantial volumetric benefits. The new configuration resulted in an increase in the integral volume of -40% for the front section of the fuselage, and an increase in the integral volume -18.5 % for the full fuselage section, and <2% L / D performance impact for a higher volume, with negligible impact to the final range. In addition, this configuration provided a larger volume for the system without impacting aircraft performance. Overall, this approach provides a26IPTS / 2OOO81736.1Attorney Docket No. RUNE-001 WO larger volume (>50% nose section volume increase vs. baseline) for cargo hold, hence increasing customer value. This balance of added volume with minimal aerodynamic cost is crucial for system packaging, such as housing avionics, cargo, sensors, or propulsion subsystems in advanced aircraft.

[0124] Ultimately, this study exemplifies how aerodynamic and structural design can be co-optimized to support expanded functionality, here, in the form of increased cargo capacity and internal system space, without sacrificing overall performance, thus delivering greater customer and mission value.A Specific Example

[0125] This section presents a concrete example of the aircraft design developed through the MDAO and system integration efforts previously discussed. The described configuration showcases a flexible and scalable platform that balances payload, range, and propulsion innovation to meet modern cargo transport needs.

[0126] The designed aircraft is able to carry a nominal payload of 1,000-4,000 pounds, which is significant for regional logistics missions. Notably, the payload is sized to fit within a standard shipping container, which streamlines ground logistics, handling, and intermodal compatibility, key factors in reducing operational complexity and cost. Utilizing a distributed propulsion system of eight electric motors, cutting-edge battery technology, and an efficient and high-aspect-ratio wing, the designed aircraft is capable of cruising at 150-225 knots, ideal for short- to medium-haul missions, over a baseline range of at least 300 nautical miles.

[0127] To further augment its endurance and payload-carrying capability, the aircraft was designed to incorporate multiple hybrid-electric propulsion technology as desired. In one embodiment, the aircraft was configured with a pair of electric turbogenerators, engines that generate electricity while burning fuel. These generators extend the aircraft’s range significantly to 400-1000 nautical miles, while also doubling its payload capability. This hybrid mode addresses one of the primary limitations of all-electric aircraft, range and payload constraints, by offering scalable energy supplementation without a complete dependence on battery storage.

[0128] The aircraft is intended to operate without onboard pilots, relying instead on a suite of onboard sensors, cameras, and autonomous navigation systems. The designed aircraft were thus equipped with onboard sensors, cameras, and navigation systems that allow the aircraft to fly and navigate without the need for human intervention. These systems enable27IPTS / 2OOO81736.1Attorney Docket No. RUNE-001 WO full autonomous flight and decision-making, removing the need for human intervention in real-time and freeing the aircraft from constraints like cockpit visibility, manual controls, or human-centered interior layouts. The unmanned configuration provides notable advantages. Without the need for a cockpit, seating, or pilot interfaces, more internal volume can be allocated to payload space, power systems, or aerodynamic optimization. Additionally, it eliminates human ergonomic limitations such as cabin pressurization, escape systems, and visibility requirements, simplifying the airframe and potentially reducing weight and cost.

[0129] In the following, specific aircraft design requirements, the conceptual design configuration selection and evaluation, energy and thrust requirements, performance characteristics, sensitivity analyses, and the CAD model showing the internal and external aircraft layout are further described in detail. An updated bill of materials is also described.

[0130] Concept of Operations

[0131] This section introduces the concept of operations (ConOps) and design compliance strategy for the aircraft, detailing both regulatory considerations and operational logistics that influence the engineering process. It lays the foundation for how the aircraft is expected to operate within real-world constraints, from certification to energy management and mission execution. The mission profile 700 for the designed aircraft is shown in FIG. 7.

[0132] Specifically, the design process was shaped by a combination of internal performance goals and anticipated regulatory requirements, particularly from the FAA’s 14 CFR Part 23. These regulations establish airworthiness standards for small aircraft (under 19,000 pounds), covering areas such as structural integrity, propulsion, flight characteristics, and system safety. While not explicitly listed here, these standards often include crashworthiness, redundancy in control systems, and minimum performance metrics for flight conditions like stall speed and climb rate.

[0133] One regulation highlighted is CFR 91.167, which specifies fuel reserve requirements for flights operating under instrument flight rules (IFR). This regulation mandates that an aircraft must be capable of flying for an additional 45 minutes at normal cruise speed after reaching its intended destination. Although this may not be strictly enforced under certain exceptions, it remains a guiding principle for sizing and energy reserve planning, particularly for electric or hybrid-electric aircraft where energy constraints are more acute.28IPTS / 2OOO81736.1Attorney Docket No. RUNE-001 WO

[0134] Another important operational aspect is battery lifecycle management. Over time, batteries degrade and lose capacity. The aircraft system design includes provisions for battery replacement once a pack has exceeded its usable life, ensuring continued performance and reliability. This also implies tracking systems for monitoring battery health and performance over time, which are vital for maintenance planning and certification compliance. For such reasons, the designed aircraft needs to fly its routes with battery charges kept at least 20% of capacity. When the battery levels are too low to meet the requirements for the next flight, the batteries need to be charged while in place on the aircraft. Batteries eventually lose their ability to retain sufficient charge and must be replaced when they have exceeded their useful life.

[0135] For sizing analysis, several methods and iterations were used to size the designed aircraft. Aircraft with similar missions, configurations, and payload requirements were used to begin the sizing. This was followed by an analysis of the specific requirements for the designed aircraft.

[0136] Together, these elements form a comprehensive operational concept that addresses real-world constraints, mission reliability, energy logistics, and certification pathways, ensuring the aircraft is not just technically sound, but also viable and compliant in practical deployment scenarios.

[0137] Sizing Analysis

[0138] A detailed explanation of the sizing analysis used to configure the electric and hybrid aircraft designs. It outlines how the aircraft was dimensioned and weighed based on performance requirements, regulatory constraints, and component specifications. The goal was to achieve an efficient design that meets mission needs, integrates standard cargo systems, and respects the limits imposed by electric propulsion technologies.

[0139] Initial sizing was conducted using the initial requirements set at the beginning of the design process. The ULD shipping container, which is commonly used in air transit operations and is the required cargo container for the designed aircraft, was used to determine the initial dimensions of the fuselage. Since the aircraft is cargo-focused, it was essential to build the airframe around standardized shipping equipment for compatibility with existing logistics systems.

[0140] Another aspect taken into consideration was the required battery volume, which is much larger than the typical volume needed for fuel. Besides this, a list was created which29IPTS / 2OOO81736.1Attorney Docket No. RUNE-001 WO included the structure of the aircraft and various components and systems within the aircraft. To get an estimated weight of each item, an internal weight model along with Raymer’s equations from “Aircraft Design: A Conceptual Approach” were used. These equations are based on historical data of many types of aircraft; the specific equations used were those focused on the general aviation aircraft since the designed aircraft would fall under FAA CFR Part 23 regulations. Notably, systems that are required for pressurization, air conditioning, and auxiliary power units are not needed for the designed aircraft, reflecting the aircraft’s unmanned, short-range nature. Table 1 below lists some weights for electric and hybrid components.Table 1 : Itemized weights for electric and hybrid configurations(*Batteries not included and weights for certain components are not listed)

[0141] The battery was sized separately, to meet the power requirements as determined by the constraint analysis (mentioned below). With these requirements, the energy needed for each leg (or mission phase) of flight was calculated based on the constraint analysis and used to determine the battery weight to meet these needs. The batteries used for this weight analysis were commercial pouch cell batteries which are listed as having a capacity of 6,604 milli-Amp hours, a nominal voltage of 3.45 volts, and a weight 46.36 grams per cell. These cell batteries may have an energy density of at least 250 Wh / kg. A factor of 1.5 was multiplied by the calculated required battery weight to account for the structure needed for battery casing and cooling the battery.

[0142] The time for the take-off and climb segments was calculated based on the constraint analysis and rate of climb, while the cruise segment endurance was determined by how much weight allowance was left in the aircraft for the batteries. Subtracting the weight of all other required components from the MTOW resulted in the allowable battery pack30IPTS / 2OOO81736.1Attorney Docket No. RUNE-001 WO weight. Note this method does not consider reserve energy allowances, typically 20% to maximize battery life.

[0143] Once the calculations were completed for the full electric and hybrid aircraft, it was found that the aircraft’s empty weight fraction was consistent with other aircraft within this size and mission class.

[0144] In the next, a constraint analysis was conducted to determine the optimum T / W ratio and W / S for given flight conditions and requirements. This method is similar to available public methods in Raymer, Gudmundsson, etc. The T / W was calculated for a given wing loading at each constraint, and the required power was calculated by converting the thrust requirement into BHP. Estimations were made for values not known at the beginning stage of the design process, and the constraint analysis was updated to reflect changes during the design process. Requirements set by the initial design parameters were also implemented. The constraint diagrams are represented by FIGS. 8A-8B.

[0145] Incorporating the concept of augmented lift, where propellers on the leading edge of the wing and large flaps at the trailing edge create greater lift at lower speeds, a design point was chosen to maximize wing loading while keeping the wing area as small as possible. The most stringent constraints for the design are the rate of climb and the stall speed. Note that the stall speed is not a hard limit as the CFR Part 23 requirements simply state that the stall speed must be reasonable such that the average remote operator can control the aircraft for a safe landing. The design point was selected at a wing loading of about 50 Iblftl, which is relatively high for aircraft of this class but justified by the enhanced lift mechanisms. This smaller wing area helps reduce structural weight, drag, and possibly cost, while maintaining adequate aerodynamic performance. The T / W ratio was chosen to be approximately 0.16, based on a balance between climb performance, propulsion system size, and energy efficiency. From FIG. 6B, at the chosen wing loading and T / W ratio, 950 BHP is the maximum horsepower required to maintain flight at all constraint conditions. This value drives the sizing of the propulsion system and energy storage (for electric and hybrid-electric configurations), ensuring the aircraft can consistently operate across climb, cruise, and maneuvering scenarios.

[0146] Performance

[0147] The constraint diagrams as described above were used primarily to determine the aircraft’s ability to meet the performance requirements, while the energy analysis of the31IPTS / 2OOO81736.1Attorney Docket No. RUNE-001 WO batteries was used to determine endurance and range. This section outlines the aircraft’s takeoff performance, emphasizing how blown lift, engine power, and altitude interact to influence required runway lengths and energy demand. It provides both analytical and numerical insights into how the design meets FAA CFR Part 23 takeoff performance standards, which include clearing a 50-foot obstacle after takeoff.

[0148] For take-off performance, the ground run distance was determined based on the kinematic relations between acceleration, speed, and distance:SG = S - So =v2~vS)(1) where a is the average acceleration during take-off calculated at The total requireddistance for takeoff is the sum of the ground run (where the aircraft accelerates on the runway), rotation (the transition from ground roll to liftoff), transition (the lift-off arc into the climb), and climb to fly over a 50-foot obstacle, as mandated by regulations under CFR Part 23 requirements.

[0149] Using the constraint diagram which accounts for blown lift, 950 hp gives a 2500 ft takeoff distance at 10,000 ft. Only 733 hp is required to achieve a 2500 ft takeoff distance at sea level. These statements show how sensitive the horsepower requirement is to takeoff altitude.

[0150] Configuration

[0151] Many configurations were considered for the designed aircraft, including a blended wing-body, a twin-boom similar to the Cessna Sky Courier or P-38, a high aspect ratio truss-braced wing, and a cantilevered high-wing aircraft. Each of these possible configurations was evaluated based on the design priorities, including a high lift-to-drag ratio, cargo loading accessibility, battery accessibility, and structural efficiency with low overall aircraft weight. Table 2 shows each of the configurations with their advantages and disadvantages.Table 2: Pros and cons of various aircraft configurations32IPTS / 2OOO81736.1Attorney Docket No. RUNE-001 WO

[0152] The eventual selection of the aircraft design was to combine the concepts of the high aspect ratio wing while reducing the planform area to reduce the wingspan. The smaller wingspan can accommodate safer ground operations and the elimination of the strut would give clearance for a large cargo door. The blown lift effect due to the distributed propulsion system along the leading edge of the wing increases the coefficient of lift (e.g., by at least 40% during takeoff), allowing for a higher wing loading, reducing the required wing area and associated weight. Here, the blown lift effect refers to the increase in aerodynamic lift generated when high-speed airflow, usually produced by propellers, fans, or jets, is directed over an aircraft’s wing surfaces, especially at low speeds. This phenomenon is particularly valuable during takeoff, landing, and slow flight, where generating sufficient lift is critical.

[0153] After completing some initial sizing calculations, it was discovered that the initial MTOW goal would not be sufficient for the desired range with the required payload and batteries. After iterative sizing, a higher MTOW was selected, still within FAA Part 23 limits, which improved both range and payload capacity without breaching regulatory thresholds.

[0154] In the first iteration of the design, the airline class container was chosen over the non-standard container size. FIG. 9A shows an early CATIA model 900 of aircraft showing the placement of a standard ULD container. FIG. 9B shows the side, front, and top views 950 of the designed craft as analyzed here.

[0155] The high wing has a reduced planform area and increased aspect ratio to reduce weight and maximize lift. The distributed electric propulsion system was made up of a total of eight 5-bladed propellers positioned along the leading edge of the wing. The generous spacing between the fuselage and inner propellers allows clearance for cargo loading, the distance between the outboard propeller and wingtip was selected to ensure reduced risk of prop and motor damage during ground operations, and the spacing between prop tips was optimized to maximize lift effects. The sleek composite fuselage accommodated the ULD33IPTS / 2OOO81736.1Attorney Docket No. RUNE-001 WO container while minimizing the cross-sectional area, and the landing gear was kept short to minimize drag.

[0156] FIG. 10 shows a later iteration 1000 of the designed aircraft with a slightly larger fuselage to accommodate multiple ULD containers for the hybrid version, a T-tail to move the horizontal stabilizer out of more turbulent flow, and a slight wing taper along the aft edge (i.e., trailing edge) of the wing with an increased root chord and decreased tip chord without changing the overall wing area, maintaining aerodynamic efficiency and control.

[0157] The internal layout of the electric or hybrid-electric versions is similar to one of them shown in FIG. 11 (which shows an internal layout of an aircraft), where turbogenerators and batteries are placed such that it meets the overall aircraft center of gravity (CG) requirement. The small amount of fuel needed for the generators is located in small tanks on either wing.

[0158] It was found that the fuselage had excess volume, so a later iteration involved a slight fuselage enlargement to accommodate an additional ULD container. The two containers were positioned at the CG of the aircraft with a cargo door aligned with one of the containers and a floor that would allow for sliding the second container into position.

[0159] Aerodynamics

[0160] As an electric aircraft, the designed aircraft must have optimal aerodynamics to perform as needed in flight. A careful examination of low drag and high lift characteristics will minimize the energy lost to maximize its range. Due to the configuration selection, three airfoils (an airfoil refers to the cross-sectional shape of a wing, blade, or sail (basically any surface that moves through air and generates lift)) were selected for the wing, vertical tail, and horizontal tail. The tail airfoils were selected based on control characteristics, and the wing required excellent lift-to-drag performance. High lift devices are designed for improved performance during critical phases of flight; flaps and blown lift will be used to generate high amounts of lift during the takeoff and landing phases of flight. Lift and drag models were analyzed using hand methods and various open-sourced as well as commercial software.

[0161] The configuration of the designed aircraft requires airfoils for the wing, horizontal tail, and vertical tail. The tail surface airfoils were chosen based on control characteristics: a symmetric NACA airfoil for the vertical tail and horizontal tail. The aircraft wing sizing study resulted in a wing area of over 200 ft2and aspect ratio of 15-18 with a design requirement of Vstaii = 58 KEAS. To meet this goal, an aircraft CLmax of >4 was required.34IPTS / 2OOO81736.1Attorney Docket No. RUNE-001 WOThe design cruise condition is 150-200 KTAS at 10,000 ft. In order to achieve this level of performance, the required airfoil was chosen, the same airfoil that was developed for use on other similar projects from NASA.. This low-drag high-Cl airfoil has a minimum drag coefficient of 0.0053 at Cl =0.90 and achieves laminar flow back to a large area of the wing upper and lower surface.

[0162] The aircraft was limited to a reasonable wingspan (under 80 ft) to minimize the risk of damage with ground operations. To achieve maximum efficiency, an aspect ratio of 15-17 was targeted to achieve the required range. The wing was designed with no taper to reduce manufacturing costs and to ensure the chord was large enough to accommodate the required systems in the wing.

[0163] The high-lift system consists of 4 propellers and nacelles (i.e., motor casings) distributed along the leading edge of each side of the wing in addition to trailing edge flaps which provide not only forward propulsion, but also provide the effect of augmented lift. The effect of the increased airflow over the wing from the distributed electric propulsion gave a calculated increase in the coefficient of lift (CL) of 73%.

[0164] Table 3 features a summary of the aerodynamic effects of the inclusion of high-lift devices. This table also serves as an estimation of the takeoff and landing configuration 3D maximum coefficient of lift (CL, Max) values. The clean wing would be representative of a cruise configuration, while the high-lift system represents the takeoff and landing configuration.Table 3: Effect of high-lift devices

[0165] Multiple methods were used to model the lift and drag of the aircraft to ensure that reasonable values were obtained. VSPAero was used to estimate properties based on a meshed 3D geometry of the aircraft using both vortex lattice and panel methods. Traditional methods using DATCOM and methods defined in Raymer’s text were also used for comparison.35IPTS / 2OOO81736.1Attorney Docket No. RUNE-001 WO

[0166] VSPAero was used to gain preliminary estimates of how the main components worked together to generate lift and drag throughout the flight envelope. The model used for analysis assumed a smaller cross-sectional area than the CAD model details. In order to account for the increase in drag due to a larger cross-sectional area, an equation was derived from DATCOM 4.2.3.1 for the sensitivity of the body drag to changes in the diameter of the body. This effect is generally nonlinear. However, when showing the sensitivity, a linear relationship can be assumed for small differences in diameter. Different speeds, angles of attack, and angles of sideslip were tested over many iterations.

[0167] By using a blend of VSPAero simulations, empirical DATCOM adjustments, and Raymer-based calculations, the disclosure achieved a robust, multi-validated aerodynamic model. This not only confirmed the aircraft’s performance targets but also provided confidence that lift, drag, and stability characteristics would remain reliable across flight conditions and design refinements.

[0168] Propulsion

[0169] This comprehensive section details the propulsion system of the designed electric and hybrid-electric aircraft, covering motor selection, distributed propeller layout, battery specifications, and the proposed hybrid augmentation to address performance limitations of a purely electric setup.

[0170] Electric motors are more sustainable and require less maintenance than traditional propulsion options. Out of a list of several commercially available motors as shown in Table 4, an appropriate motor was selected to provide required range, maneuverability and performance.Table 4: Electric motor comparison36IPTS / 2OOO81736.1Attorney Docket No. RUNE-001 WO

[0171] Specifically, an electric motor with an advanced inverter integrated within the motor system was selected. The inverter is an essential component responsible for converting the direct current (DC) from the aircraft’s battery into the alternating current (AC) required to power the motor. It streamlines the aircraft’ s powertrain architecture, reducing the need for additional external components and wiring, thereby saving space and weight. The integrated inverter ensures seamless power delivery and enhances the motor’s efficiency by optimizing the conversion process. This results in reduced power losses and greater utilization of the available energy, leading to longer flight times and extended range capabilities.

[0172] For propeller selection and sizing, the designed aircraft makes use of an efficient distributed electric propeller configuration. The location and size of the propellers were driven by the wingspan, fuselage width, and cargo door accessibility. With a set wingspan of under 80 ft and fuselage width of under 7 ft, the available length for the propellers for each wing was computed. When choosing the number of propellers to distribute across the leading edge of this high-wing aircraft, it is important to consider the space needed to provide clearance for a cargo door in the fuselage. Additionally, keeping the outboard propeller fully inboard of the wingtip was prioritized to reduce the risk of damage during ground operations.

[0173] Considering these points, an 8-propeller configuration was selected, which brings numerous benefits while adhering to all aircraft limitations. The first advantage is the provision of redundancy in case of a failure of one or more propellers. This redundancy significantly enhances the safety of the aircraft, as it can maintain stable flight even if some propellers are not functioning optimally. The second advantage is the improved efficiency achieved through distributed electric propulsion. This setup allows for better power distribution and control, resulting in increased overall efficiency. By adjusting the thrust of individual propellers, the aircraft can optimize its performance for different flight phases, including takeoff, cruising, and landing, leading to fuel savings and extended flight ranges.

[0174] A total of eight five-bladed propellers were chosen for the distributed propulsion system. The required propeller diameter was calculated based on the power required using Gudmundsson’s equation 15- 29:Dp = Kp PBHP25(2)IPTS / 2OOO81736.1Attorney Docket No. RUNE-001 WO

[0175] The propellers were evenly distributed along the leading edge of the wing with a spacing between propeller tips to take maximum advantage of the lift effects. Distributed electric propulsion can help reduce aircraft noise levels by using smaller and more numerous propellers driven by electric motors versus traditional internal combustion engines. Optimization of propeller placement for performance, blown lift effects, and noise reduction were further investigated using CFD studies.

[0176] The propellers are expected to be made of composite materials, such as carbon fiber or fiberglass, providing a high strength-to-weight ratio, resulting in lightweight yet durable propellers. The 5-blade configuration enhances aerodynamic efficiency, reducing noise and vibrations while improving thrust distribution and overall performance.Additionally, the increased number of blades allows for better load distribution, enabling the propeller to handle higher power outputs more effectively. CFD simulations provide crucial insights into the propeller’s performance and efficiency, ensuring that the chosen Kp value is appropriate for the specific design and contributing to the optimal functioning of the aircraft’s propulsion system.

[0177] For a variable pitch propeller, the altitude has little effect on its efficiency due to the nature of its design and functionality. The propeller pitch can be automatically adjusted to optimize performance at various altitudes during flight, enabling the aircraft to maintain the desired motor speed and maximize efficiency.

[0178] The maximum motor revolutions per minute (RPM) was used for thrust calculations in the disclosed distributed electric propulsion system. The decision to use the maximum RPM was primarily driven by the need to assess the propellers’ maximum potential thrust output, which is crucial for understanding the overall performance of the aircraft during takeoff and climb phases. By employing the maximum RPM values, the highest possible thrust levels that the propellers could generate can be analyzed, providing valuable insights into the aircraft’s acceleration and initial climb performance.

[0179] It is to be noted that as this technology progresses and further studies are conducted, the RPM used for different flight phases may be subject to modification. The use of the maximum RPM might not be the most efficient option during all stages of flight. For instance, during the cruise flight, maintaining the propellers at their peak RPM might lead to increased power consumption and, consequently, reduced overall efficiency.38IPTS / 2OOO81736.1Attorney Docket No. RUNE-001 WO

[0180] One of the crucial considerations that prompted to use the maximum RPM during thrust calculations was to ensure that the angular velocity at the tip of the propellers would not reach excessively high values, such as Mach 0.8. Operating the propeller tips at such high speeds could result in compressibility effects and undesirable aerodynamic phenomena, which could compromise the aircraft’s performance and structural integrity. By employing the maximum RPM value and carefully monitoring the tip speeds, it can be ensured that the propellers operate within safe and optimal ranges, minimizing the risk of adverse effects associated with supersonic blade speeds.

[0181] For battery analysis, several batteries were considered for the designed aircraft. The main requirement was a high energy density. Battery test information was used to gain a realistic expectation of new and upcoming batteries.

[0182] For the battery analysis phase, a system of equations was used to find the number of cells, the weight of the pack, and the volume of the battery pack. To initially benchmark the batteries, an 800-volt motor at 1000 horsepower was used for 1 hour to find which system would weigh the least and include the fewest number of cells for those energy needs. Table 5 below is an example list of select commercial batteries considered during this project.Table 5: List of considered batteries39IPTS / 2OOO81736.1Attorney Docket No. RUNE-001 WO

[0183] The battery from above table, providing the low weight, required discharge rate and compact packaging was selected. As battery technology improves, resulting in higher energy densities, the weight of the batteries should only decrease as compared to the batteries analyzed here.

[0184] While studying the battery system, it was noted that the capacity was the main driving force. Though a higher nominal voltage would still be beneficial, a larger capacity would be even more important. It was also noted that with a larger capacity comes a greater weight in the battery. Table 6 summarizes the battery system requirements for the given aircraft configuration.Table 6: Select battery system requirements

[0185] The electric version of the aircraft faces significant limitations in range, speed, and payload capacity. To address this, a hybrid-electric configuration using the same airframe was proposed. The selected powerplant operates on conventional and sustainable aviation fuels to generate electricity for hybrid propulsion.

[0186] FIG. 12 shows some of the sizing capabilities of the hybrid configuration, such as payload (y) vs. range (x), as defined by equations in the figure, according to one embodiment.

[0187] Stability and Control

[0188] This section focuses on the stability and control characteristics of the aircraft, including the empennage configuration, and both longitudinal- and lateral-directional stability assessments, using a combination of analytical and simulation tools.40IPTS / 2OOO81736.1Attorney Docket No. RUNE-001 WO

[0189] For empennage, the tail configuration chosen for this design is a conventional tail, which means a horizontal stabilizer mounted on the lower rear fuselage, and a vertical stabilizer above it. The conventional tail was chosen for its low structural weight, simplicity in comparison to a T-tail or V-tail, and ease of control surface actuation. The trade-off is its potential exposure to propwash, particularly from the distributed electric motors mounted along the leading edge of the wing. The position of the horizontal tail is critical to ensure that the airplane is controllable throughout the various stages of flight. The airfoils of both the vertical and horizontal tails were thin, symmetric airfoils.

[0190] The longitudinal stability (i.e., the aircraft’s ability to maintain or return to level flight after a disturbance in pitch) was evaluated using the same methods as for the lift and drag models: DATCOM and VSP. Both methods showed the aircraft as stable, DATCOM more than VSP, however not all effects were taken into account. For the DATCOM methods, only the effects of the wing, body, and tail were considered along with their interferences with each other. The VSP panel method analyzed the wing, tail, and body including the interferences with each other. The model that was used to calculate the stability values assumed a smaller cross-sectional area for the fuselage and only one cargo container inside the fuselage. Despite these gaps in the stability model, both methods showed plenty of margin (shown in FIG. 13) for destabilizing effects. FIG. 14 also shows evidence of longitudinal stability.

[0191] With respect to the lateral and directional stability, the VSP panel method was solely utilized to analyze the model. The results were initially unstable in both cases. However, the tail was iteratively resized to reach stability for lateral, directional, and yaw stability. The analysis was performed at a cruise speed at a cruising altitude of 10,000 ft over a range of angles of sideslip from -10 to 10 degrees. FIGS. 15A-15C show the stability of the aircraft graphically.

[0192] From the above, it can be seen that the aircraft’s final configuration achieves longitudinal stability (such as pitch control and center of gravity management) and lateral and directional stability (such as roll and yaw control, especially during turns and crosswind conditions). These analyses confirm that with proper tail sizing and configuration, the aircraft remains controllable, certifiable under FAA Part 23 standards, and robust across varied mission conditions, despite the complexities introduced by the distributed propulsion system and blown-lift aerodynamics.41IPTS / 2OOO81736.1Attorney Docket No. RUNE-001 WO

[0193] Mass Properties

[0194] In general, the designed aircraft uses the lightest weight materials with the smoothest finishes to minimize weight and drag.

[0195] With respect to the major component weights and locations, the center of gravity of the aircraft was determined by estimating the location of individual aircraft component items of mass using the internal data as well as available public domain information from references such as Raymer. Individual item locations affect the aircraft center of gravity and, therefore, the stability. Taking care to place the items of mass at locations that result in a stable aircraft was a priority in this exercise.

[0196] With respect to the center of gravity envelope, the center of gravity should remain constant throughout the flight as much as possible. The payload location is very close to the aircraft center of gravity so that when the aircraft is unloaded, the center of gravity moves forward only slightly.

[0197] For the hybrid aircraft, the center of gravity was designed to be very close to that for the electric-only aircraft by placing the fuel in the wing so that it aligns with the payload FS. With respect to the determination of moments and products of inertia, the moments of inertia are necessary to determine the dynamic stability of the aircraft. They describe the aircraft’s resistance to change in its rotational rate about any axis. The designed aircraft’s moments of inertia about the X, Y, and Z axis (Ixx, Iyy, Izz) were determined using a method from Raymer’s textbook.

[0198] Trade Studies

[0199] This section addresses the mass properties of the aircraft, and how weight distribution affects center of gravity (CG), stability, and dynamic behavior, with particular focus on component placement and the aircraft’s resistance to rotational motion.

[0200] Several trade studies were conducted throughout the design. The wing loading and thrust-to-weight ratio were determined based on the constraint diagram’s illustration of acceptable combinations that would meet the design goals. The aircraft lift coefficient needed to achieve the climb and takeoff requirements was determined after adjusting the aspect ratio, wing area, and maximum takeoff weight.

[0201] FIG. 16A shows the importance of maximizing the lift-to-drag ratio. FIG. 16B shows the speed vs. range, which led to the reduction in required speed, a non-issue for short-42IPTS / 2OOO81736.1Attorney Docket No. RUNE-001 WO range flights. FIG. 16C illustrates where the designed aircraft lies related to other electric and hybrid aircraft in terms of speed and range.

[0202] When determining the best maximum takeoff weight for the design, the sensitivity of MTOW on range and payload was studied. Once the required MTOW was selected, a customer-facing range vs. payload sensitivity was studied, the result of which is shown in FIG. 17.

[0203] For the hybrid-electric version, trade studies were performed to show the range, payload, and cruise speed for one vs. multiple turbogenerators.

[0204] The aircraft as designed would be able to accommodate this amount of fuel for the selected number(s) of turbogenerators.

[0205] Finally, the drag vs. cross-sectional area was investigated to better understand the consequences of any growth in the fuselage that may be necessary as the design progresses FIG. 18 shows the effect of fuselage body diameter on aircraft drag.

[0206] Through iterative trade studies and constraint diagram analysis, the most efficient and certifiable combination of wing loading, thrust-to-weight ratio, and lift capability was determined. This allowed the aircraft to meet its performance targets, particularly for short takeoff, steep climbs, and energy-efficient cruise flight, while also supporting the needs of both electric and hybrid propulsion architectures.

[0207] Bill of Materials

[0208] This section summarizes the bill of materials (BOM) considerations and ties together the major design elements and trade studies that shaped the final configuration of the electric and hybrid aircraft.

[0209] The electric and hybrid aircraft would look similar on the outside. The internal structural supports may be slightly different to accommodate the different power systems. The detailed BOM was developed based on the internal database as well as historical cost.

[0210] The preliminary design of the disclosed autonomous electric cargo aircraft was completed based on the design priorities. The initial configuration of a high-aspect ratio high- wing aircraft with a conventional tail and short landing gear accommodates at least one ULD cargo container and is propelled by eight electric motors and propellers distributed along the leading edge of the wing. The ULD container can be accessed through a large, reinforced door on the side of the fuselage.43IPTS / 2OOO81736.1Attorney Docket No. RUNE-001 WO

[0211] Several trade studies were conducted throughout the design. The wing loading and thrust-to-weight ratio were determined based on the constraint diagram’s illustration of acceptable combinations that would meet the design goals. The aircraft lift coefficient needed to achieve the climb and takeoff requirements was determined after adjusting the aspect ratio, wing area, and maximum takeoff weight. The speed vs. range was investigated, which led to the optimum speed for the flight range. Father, the drag vs. cross-sectional area was investigated to better understand the consequences of any growth in the fuselage that may be necessary as the design progresses.

[0212] The performance of the designed aircraft meets the goals through its use of some of the highest energy density batteries currently available and enhanced lift from the distributed electric propulsion. As battery energy density improves, the aircraft range, speed, and payload capabilities also improve. To meet projected demands, a hybrid electric version of the aircraft was developed. This hybrid version, since it is capable of carrying higher payload with higher speed, can accommodate multiple ULD containers.System and / or Computer Embodiments

[0213] FIG. 19 depicts an example computing device 1900 for implementing systems and methods described in reference to FIGS. 1-18. Examples of a computing device may include a personal computer, desktop computer laptop, server computer, a computing node within a cluster, message processors, hand-held devices, multi-processor systems, microprocessorbased or programmable consumer electronics, network PCs, minicomputers, mainframe computers, mobile telephones, PDAs, tablets, pagers, routers, switches, and the like.

[0214] In some embodiments, the computing device 1900 includes at least one processor 1902 coupled to a chipset 1904. The chipset 1904 includes a memory controller hub 1920 and an input / output (I / O) controller hub 1922. A memory 1906 and a graphics adapter 1912 are coupled to the memory controller hub 1920, and a display 1918 is coupled to the graphics adapter 1912. A storage device 1908, an input interface 1914, and network adapter 1916 are coupled to the VO controller hub 1922. Other embodiments of the computing device 1900 have different architectures.

[0215] The storage device 1908 is a non-transitory computer-readable storage medium such as a hard drive, compact disk read-only memory (CD-ROM), DVD, or a solid-state memory device. The memory 1906 holds instructions and data used by the processor 1902. The input interface 1914 is a touch-screen interface, a mouse, track ball, or other type of44IPTS / 2OOO81736.1Attorney Docket No. RUNE-001 WO input interface, a keyboard, or some combination thereof, and is used to input data into the computing device 1900. In some embodiments, the computing device 1900 may be configured to receive input (e.g., commands) from the input interface 1914 via gestures from the user. The graphics adapter 1912 displays images and other information on the display 1918. The network adapter 1916 couples the computing device 1900 to one or more computer networks.

[0216] The computing device 1900 is adapted to execute computer program modules for providing functionality described herein. As used herein, the term “module” refers to computer program logic used to provide the specified functionality. Thus, a module may be implemented in hardware, firmware, and / or software. In one embodiment, program modules are stored on the storage device 1908, loaded into the memory 1906, and executed by the processor 1902.

[0217] The types of computing devices 1900 may vary from the embodiments described herein. For example, the computing device 1900 may lack some of the components described above, such as graphics adapters 1912, input interface 1914, and displays 1918. In some embodiments, a computing device 1900 may include a processor 1902 for executing instructions stored on a memory 1906.

[0218] The methods disclosed herein may be implemented in hardware or software, or a combination of both. In one embodiment, a non-transitory machine-readable storage medium, such as one described above, is provided, the medium comprising a data storage material encoded with machine readable data which, when using a machine programmed with instructions for using said data, is capable of displaying any of the datasets and execution and results of this disclosure. Such data may be used for a variety of purposes, such as patient monitoring, treatment considerations, and the like. Embodiments of the methods described above may be implemented in computer programs executing on programmable computers, comprising a processor, a data storage system (including volatile and non-volatile memory and / or storage elements), a graphics adapter, an input interface, a network adapter, at least one input device, and at least one output device. A display is coupled to the graphics adapter. Program code is applied to input data to perform the functions described above and generate output information. The output information is applied to one or more output devices, in a known fashion. The computer may be, for example, a personal computer, microcomputer, or workstation of conventional design.45IPTS / 2OOO81736.1Attorney Docket No. RUNE-001 WO

[0219] Each program may be implemented in a high-level procedural or object-oriented programming language to communicate with a computer system. However, the programs may be implemented in assembly or machine language, if desired. In any case, the language may be a compiled or interpreted language. Each such computer program is preferably stored on a storage media or device (e.g., ROM or magnetic diskette) readable by a general or special purpose programmable computer, for configuring and operating the computer when the storage media or device is read by the computer to perform the procedures described herein. The system may also be considered to be implemented as a computer-readable storage medium, configured with a computer program, where the storage medium so configured causes a computer to operate in a specific and predefined manner to perform the functions described herein.

[0220] The databases thereof may be provided in a variety of media to facilitate their use.The databases of the present disclosure may be recorded on computer readable media, e.g., any medium that may be read and accessed directly by a computer. Such media include, but are not limited to: magnetic storage media, such as floppy discs, hard disc storage medium, and magnetic tape; optical storage media such as CD-ROM; electrical storage media such as RAM and ROM; and hybrids of these categories such as magnetic / optical storage media. One of skill in the art may readily appreciate how any of the presently known computer readable mediums may be used to create a manufacture comprising a recording of the present database information. "Recorded" refers to a process for storing information on computer readable medium, using any such methods as known in the art. Any convenient data storage structure may be chosen, based on the means used to access the stored information. A variety of data processor programs and formats may be used for storage, e.g., word processing text file, database format, etc.46IPTS / 2OOO81736.1

Claims

Attorney Docket No. RUNE-001 WOWHAT IS CLAIMED IS:

1. A hybrid-electric autonomous aircraft, comprising: a fixed-wing airframe sized to carry at least one standard unit load device (ULD); a distributed electric propulsion (DEP) system including one or more electric propellers disposed along a leading edge of a wing; and a hybrid power system including at least one turbogenerator and one or more battery packs to drive the one or more electric propellers.

2. The aircraft of claim 1, wherein the at least one turbogenerator is sized by a power requirement during a cruise phase, and the one or more batteries provide a required boost power during one or more of a takeoff phase and a climb phase.

3. The aircraft of claim 1, wherein the DEP system includes a number of multi- bladed propellers, each driven by an electric motor.

4. The aircraft of claim 1, wherein a battery pack includes multiple battery cells with an energy density of at least 250 Wh / kg.

5. The aircraft of claim 1, wherein the one or more electric propellers provide a augmented lift effect that increases a coefficient of lift by at least 40% during takeoff.

6. The aircraft of claim 1, wherein the aircraft is configured to have a maximum takeoff weight (MTOW) of at least 6,000 pounds.

7. The aircraft of claim 1, wherein the at least one turbogenerator operates on sustainable aviation fuel (SAF) or diesel and provides power redundancy for extended range.

8. The aircraft of claim 1, further comprising a composite fuselage configured to minimize cross-sectional area and house at least two cargo containers aligned with a center of gravity.

9. The aircraft of claim 8, wherein the fuselage includes a cargo door aligned with a floor mechanism allowing sliding of the cargo containers for loading / unloading.47IPTS / 2OOO81736.1Attorney Docket No. RUNE-001 WO10. The aircraft of claim 8, wherein the center of gravity is maintained within a flight-stable range during cargo loading and unloading operations.

11. The aircraft of claim 1, further comprising a human-in-the-loop (HITL) or human-on-the-loop (HOTL) system enabling remote operation.

12. The aircraft of claim 1, further comprising an integrated flight and propulsion control system configured to manage distributed electric propellers and control surfaces.

13. The aircraft of claim 12, wherein the control system is capable of managing both vertical and horizontal tail airfoils with symmetric NACA profiles.

14. The aircraft of claim 1, wherein the one or more battery packs are selected based on total mission energy needs.

15. The aircraft of claim 1, wherein the one or more battery packs are selected by accounting for cooling and packaging.

16. A method for designing a hybrid-electric aircraft, comprising: determining an optimal configuration of the hybrid-electric aircraft based on mission energy balance, stability, and propulsion sizing constraints by using a multi-disciplinary design analysis and optimization (MDAO) framework; implementing robust flight control laws that enable a human-in-the-loop (HITL) or human-on-the-loop (HOTL) by using a dynamics and control simulation and algorithm; verifying the HITL or HOTL through operational simulations to simulate remote operator operations; and demonstrating effectiveness of the configuration and HITL or HOTL flight control system architecture through a prototyped cost-effective subscale aircraft.

17. The method of claim 16, wherein the MDAO framework employs distributed electric propulsion (DEP) and a hybrid-electric propulsion system architecture in determining the optimal configuration of the hybrid-electric aircraft.48IPTS / 2OOO81736.1Attorney Docket No. RUNE-001 WO18. The method of claim 16, wherein the dynamics and control simulation and algorithm employs a generalized formulation of a trim problem when implementing the robust flight control laws, wherein both conventional control surfaces and propulsors act as control effectors.

19. The method of claim 16, wherein the prototyped cost-effective subscale aircraft includes a hybrid-electric subscale demonstrator that is sized using the MDAO framework to balance turbine and electric power contributions.

20. The method of claim 16, further comprising a full-scale validation, wherein a final aircraft configuration undergoes a comprehensive testing that includes ground testing, flight testing, and system-level evaluations.49IPTS / 2OOO81736.1

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