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Amphibious Aircraft Design: Conceptual Report
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Aircraft Design II - Report # 10 Juan Diego Gomez, Johanna Rodriguez, Tomas White, Andrew Mendoza, Federico Villamizar, Carlos Andres Geney, Juan Sebastian Velasquez, Emmanuel Villadiego, Harold Bejarano, Sofia Valencia, Jean Paul Sierra, Alejandro Aristizabal, Emilio Jose Guerrero, Arley Hernandez, Daniela Araque, and Lorena Pantoja Universidad Pontificia Bolivariana, Aeronautical Engineering, Medellín, Colombia I. Abstract This report presents a multidisciplinary conceptual design of an amphibious vehicle developed to satisfy the AIAA’s NEWT efficiency objectives. Starting from an assessment of mission needs, the design methodology employs studies and iterative analyses to explore and refine vehicle configurations that balance aerodynamic performance, hydrodynamic behavior, structural integrity, and weight targets. Key design drivers are identified through preliminary sizing and stability evaluations, while multidisciplinary integration ensures compatibility of propulsion, control, safety, and support systems for both waterborne and terrestrial operations. Performance predictions are generated using analytical and semi-empirical methods to verify that lift, buoyancy, drag, and load-carrying requirements are met. The design process culminates in an economic appraisal that examines cost drivers and operational considerations, thereby demonstrating the feasibility and efficiency of the proposed concept. Together, these efforts provide a cohesive overview of the conceptual-design workflow and show that the vehicle fulfills or surpasses NEWT criteria for versatility, safety, and operational economy. II. Introduction In regional contexts with limited transportation infrastructure an amphibious aircraft could significantly enhance connectivity and socioeconomic development. These aircraft offer a practical solution to overcome geographic barriers like mountains and dense jungle, especially where traditional airport infrastructure is minimal or non-existent. The integration of modern technologies makes it possible to propose a versatile and efficient platform tailored to remote operations. In Latin America, Colombia presents an ideal environment for seaplane operations due to its extensive river and coastal systems. ALMA, a local company, is already pursuing regulatory approval to operate amphibious aircraft along the Caribbean coast, aiming to boost connectivity and regional tourism [2]. The ability to land on water opens new possibilities not only for regional and remote transportation, but also positions amphibious aircraft as valuable assets in sectors such as logistics, tourism, and emergency response. The global amphibious aircraft market, valued at approximately $164 million in 2021, is projected to surpass $500 million by 2031, reflecting growing demand beyond traditional missions such as search and rescue or firefighting [1]. Designing an aircraft capable of operating from both unprepared runways and bodies of water presents a unique cross-disciplinary challenge that combines aerodynamic, hydrodynamic, and structural considerations. To meet the demanding operational environment, the design must incorporate lightweight, corrosion-resistant materials within a reinforced structure, while also ensuring static and dynamic stability. Safety features such as strategically located centers of gravity and auxiliary flotation devices are essential to withstand overturns, wave impacts, and other unforeseen conditions. In addition, the aircraft must offer short takeoff and landing (STOL) capability, competitive payload and range performance, and full compliance with international airworthiness standards—all while remaining cost-effective and easy to maintain. III. Mission requirements To fulfill the expected operating role, the aircraft must comply with the following mission requirements imposed by the RFP (Request for Proposal). A. General requirements The main objective was to design an amphibious aircraft capable of transporting both passengers and/or cargo, set to enter service in 2031. To achieve this, all requirements were categorized into general and mission-specific criteria. The general requirements were as follows: 1 Requirement Goal value Aircraft type Amphibious plane Operational scenario Fresh and saltwater - improvised runways. Minimum Cruise Speed 200 knots Target Cruise Speed 250 knots or greater Flight Rules VFR and IFR capabilities Powerplant EIS in 2031 Certification rules FAR 23 (commuter type aircraft) Table 1 Mission general requirements The RFP adopts a three-mission approach that, while based on the same mission profile, set different performance goals for each mission, these being the following. B. Mission requirements for passenger mission Requirement Goal value Crew 1 Passenger capacity 19 Weight per passenger 193.6 lb Weight per passenger bag 37.4 lb Total payload weight (passengers + baggage) 4582 lb Volume per passenger bag 4 ft^3 Seat pitch 28 in Takeoff and landing field 1500 ft Range 250 nmi Table 2 Performance requirements for passenger mission C. Mission requirements for cargo mission Requirement Goal value Payload 5000 lb Range 200 nmi Operational readiness time 60 minutes Table 3 Performance requirements for cargo mission 2 D. Mission requirements for economic mission Requirement Goal value Passenger capacity 19 Weight per passenger 193.6 lb Weight per passenger bag 37.4 lb Total payload weight (passengers + baggage) 4582 lb Range 150 nmi Table 4 Performance requirements for economic mission IV. Baseline After identifying the initial RFP requirements, the next step was to establish a baseline using aircraft with similar characteristics. This approach allowed the design team to determine an optimal range of parameters suited to this category of aircraft, using the baseline as a reference for key design criteria. - DHC-3 - Sea Plane Cessna 208 Max Cruise [kts] Service Ceiling [ft] Range [nm] Takeoff Field [ft] Landing Field [ft] MTOW [lbs] Empty Weight [lbs] Payload [lbs] Passenger Capacity (with crew) Max. Continuous Power [HP] Wing Span [ft] Length [ft] Wing Area [ft²] Aspect Ratio Number of engines W/P W/S 138 18500 750 1980 1510 8000 5287 2100 186 25000 970 1160 1625 8000 4760 3532 DHC-6 Twin Otter 400 170 26700 400 1490 1510 12500 7445 3250 219 20000 810 1670 1600 14551 9250 3968 Dornier Seastar 180 15000 900 2245 2620 11240 8598 2645 13 10 22 21 11 21 21 21 9 740.24 58.07 41.99 279.86 8.97 1 10.8 28.6 867.64 1662.86 1700.41 1298.11 1528.76 3256.00 2220.73 2387.02 18548 4760 52.17 37.73 279.43 9.70 1 9.2 28.6 64.99 49.51 452.08 10.10 2 7.5 27.6 65.55 47.31 374.58 11.45 2 8.6 38.8 58.20 41.67 329.38 10.28 2 8.7 34.1 55.68 54.33 344.44 9.00 2 9.5 42.1 58.96 57.84 310.00 10.90 2 5.2 54.7 72.24 55.12 441.00 9.90 2 8.6 43.1 104.00 63.85 1399.31 7.73 2 14.8 25.2 108.76 109.78 1463.89 8.10 2 5.7 71.8 93.83 65.03 1076.39 8.20 2 9.2 40.7 L-410 Table 5 223 25000 715 2600 1480 14495 8600 1200 Beech 1900D 260 25000 708 3737 2720 16950 11000 2087 Cessna Sky Courier 210 25000 940 2740 3010 19000 12325 6000 Consolidated PBY 109 13124 219 1750 2800 35274 20910 - ShinMaywa US-2 260 23606 2300 920 1080 105160 56504 - Bombardier 415 150 14700 1319 2674 2182 43850 29983 6393 20 18 Dornier 228 Aircraft Baseline After collecting baseline aircraft data, we plotted takeoff weight against cruise speed, range, takeoff field length, and aspect ratio, fitting trend lines to each scatter plot. We then imputed the estimated takeoff weight into those trend lines to derive statistical estimates for each parameter. From this analysis, only the aspect ratio estimate, 10.5, was retained for subsequent work. 3 Fig. 1 Statistical AR graph V. Mission profile There are essentially three missions that this aircraft must fulfill—two focused on passenger transport and one on cargo operations—with variations primarily in payload and range, as specified in the RFP. Figure 2 provides an illustrative representation of the general mission profile expected across all variants, encompassing key phases such as warm-up, taxi, takeoff, climb, cruise, loiter, descent, and final landing and shutdown. Notably, none of the missions explicitly define whether the operations are to be conducted from land or water. Fig. 2 Illustrative Mission Profile VI. Conceptual design As part of the conceptual development phase, multiple preliminary sketches were generated as seen in figure 3 each exploring different configurations and operational concepts. After presenting these initial ideas, a convergence process was carried out to extract and combine their most promising elements. This led to a refined set of sketches that reflect a common direction—highlighting, for example, the adoption of a high-wing configuration, which offers clear advantages for water operations, including protection from spray and enhanced stability. The selected concepts align with the operational requirements outlined in the RFP and serve as the foundation for subsequent design iterations. 4 Fig. 3 Napkin design A. Survey To incorporate user perception into the amphibious aircraft design, a public survey was conducted to assess trust and awareness regarding this type of aircraft. Both existing and conceptual designs were evaluated, results plotted in figure 4 and 5 respectively . Data showed that the aircraft’s bow design often raises concerns, as it interacts directly with water during takeoff and landing, while floats were seen as more reassuring due to their visible stability. Among the conceptual sketches, designs that resembled conventional models (such as the Twin Otter with floats) received higher trust ratings due to their familiarity. However, float-only configurations showed operational limitations on land, especially without retractable landing gear. These findings highlight the need to balance performance with public confidence, and were considered for the final design. Fig. 4 Survey results for existing aircraft 5 Fig. 5 Survey results for conceptual aircraft VII. Weight estimation The weight fractions method was implemented to obtain the data presented below on the basis of the mission profile discussed before. However, throughout the project, as the design advanced, iteration was made to obtain values closer to the reality, reducing the uncertainty, although the data presented below was used for reference as a first approach, but the weight used for the matching chart a forward calculation was 16607 lb Phase Fraction Notation Start Taxi T.O Climb Cruise Loiter Descent Landing, taxi, shutdown 0.992 0.990 0.996 0.985 0.963 0.993 0.990 0.992 π1 /π0 π2 /π1 π3 /π2 π4 /π3 π5 /π4 π6 /π5 π7 /π6 π8 /π7 Mass fraction of fuel 0.903 Table 6 Concept Pax weight Baggage weight Payload Crew Weight Range [NM] Fuel reserves Cruise speed [kt] 3678.4 710 4389 193.6 250 1.25 250 πππππ’ππ π _ππ‘ππ ππ‘πππ 18130.3 Table 7 Fuel Fractions Value [lb] Info Pax Mission Concept Weight [lb] Concept Value Required fuel Fuel + reserves Operational Empty Trapped fuel Empty 1757.243 2196.554 11544.710 90.651 11260.459 Payload [lb] Crew Weight [lb] Range [NM] Fuel reserves [] Cruise speed [kt] 5000 193.6 250 1.25 250 Table 8 Operational and Empty Weight [lb] Table 9 6 Info Cargo Variable Value π π [h] ππ πΏ/π· πloiter [kt] πΈ [h] 0.77 0.5 15 160 0.20 Table 10 Loiter Table 12 Summary [lb] Concept Value [lb] πππ ππΉ π ππΏ ππ π. πΈπ ππ‘ π¦ Used Fuel 17534.3 1094.2 5193.6 11246.5 1519.9 Variable Value π [NM] π [kt] πΏ/π· ππ 250 250 16 0.6 Table 11 Cruise VIII. Matching chart For our mission profile, few restrictions are given, being the most restrictive of the STOL features, to ensure a design able to fulfill such requirements the matching chart approach was implemented. It was also taken into account the regulatory restrictions for aircraft of 19 passengers; it is established that such aircraft must be able to take-off with one engine inoperative under certain conditions, to ensure the performance needed, such condition was included in the matching chart. The following plot is the cleanest representation of the selected features for the aircraft. Fig. 6 Matching Chart 7 Table 13 Aircraft Design Point Parameter π [lb] πππ /π [lb/ft2 ] π π€ [ft2 ] π/π π΄π π πTot [lb] πππΈ [lb] Value 16607 27.466 604.63 0.770 12 0.75 12781.43 6390.72 IX. Initial sizing An initial dimensional approximation of the amphibious aircraft was carried out, estimating key components such as the wing, empennage, fuselage, landing gear, and a preliminary center of gravity location. This process relied on statistical methods outlined in design references by Raymer and Roskam, offering a solid foundation for early-stage sizing. To better align with the mission requirements, a comparative analysis of reference aircraft with similar roles and configurations was conducted, focusing on parameters like fuselage length and wing parameters, consigned in Table 14 and 15, respectively. These benchmarks helped guide the preliminary dimensions. It should be noted that some of these dimensions were later refined to better suit the specific operational and hydrodynamic characteristics of amphibious operations, rather than relying solely on generalized statistical trends. Fuselage sizing Length π 1, 05 π 0, 4 πΏ π [ft] 52, 33 Height and tail cone πΏπ 7 ππ πΏππ 3 ππ Θππ 8 π π [ft] 7, 48 πΏ π π [ft] 22, 43 Table 14 Fuselage sizing Regarding wing dimensions, key parameters such as the aspect ratio (AR) and taper ratio (π) were defined to optimize performance. An AR value of 10.5 was selected using a cubic regression approach, ensuring consistency between the calculated weight and the derived function from the baseline. A taper ratio of 0.7 was chosen to improve load distribution across the wingspan, enhancing aerodynamic efficiency and stability. Additionally, a leading-edge sweep angle (Λ) of 2.5° was incorporated to improve aesthetics for a better market output, particularly in cruise conditions, without compromising low-speed stability. No dihedral angle was applied, as the high-wing configuration inherently provides sufficient lateral stability. 8 Wing sizing Principal Data π π€ [ft2 ] 604, 63 Wingspan [ft] 79, 68 πΆroot [ft] 8, 93 πΆtip [ft] 6, 25 MAC (trapezoid and cte taper) length πΆbar [ft] 7, 67 Position πbar [ft] 18, 75 Flap and aileron πΆail 0, 25 πΆloc π ail 0, 45 π effective π ail [ft] 16, 25 πΆail,tip [ft] 1, 56 π flap 0, 50 π effective π flap [ft] 18, 05 Table 15 Wing initial sizing For this initial sizing, a T-tail configuration (as seen in Table 16) was considered along with engine placement near the mid-wing section. However, this layout will be revisited in later design stages to ensure optimal stability, weight distribution, and hydrodynamic performance. Empennage sizing Horizontal stabilizer Vertical stabilizer Dihedral angle [deg] 3 Dihedral angle [deg] 90 Incidence angle [deg] 0 Incidence angle [deg] 0 π΄π h 4, 5 π΄π v 1, 5 Sweep angle πΆ/4 [deg] 0 Sweep angle πΆ/4 [deg] 0 πh 0, 7 πv 0, 7 πΆele πΆele 0, 45 0, 40 πΆloc πΆloc π h [ft] 21, 53 π v [ft] 11, 73 πΆroot,β [ft] 5, 63 πΆroot,π£ [ft] 9, 20 πΆtip,β [ft] 3, 94 πΆtip,π£ [ft] 6, 44 π ele [ft] 21, 53 π rudd [ft] 11, 73 πΆele [ft] 2, 53 πΆrudd [ft] 3, 68 πΆbar,β [ft] 4, 83 πΆbar,π£ [ft] 7, 90 Table 16 Empennage Initial Sizing After obtaining the previously mentioned values, the aircraft’s initial dimensions where plotted in an Excel graph, providing an initial top view. 9 Fig. 7 Initial sizing aircraft top view X. Aircraft Configuration Given an initial sizing for the most basic and conventional configuration, in order to continue the design process it is necessary to define the configuration of the aircraft; a commonly used configuration on amphibious aircraft is the T tail with a high wing, due to the harsh environment and the adverse conditions during waterborne take off, such configuration allows to keep aerodynamic surfaces; whom are known to be thin structures, away from water. In order to maximize the aerodynamic efficiency, a V-tail was selected as the best option to ensure stability while being the option where frontal area could be minimum. Another key aspect of the configuration is location of the engines. For the analysis three possible configurations where evaluated, conventional engines on the wings like the CL-415, over the fuselage like the Dornier Seastar, and in the tail like the Beriev A-40, and the options contemplated where turbofan and turboprop. After a detailed comparison that will not be extensively discussed here, the optimal configuration was determined to be turbofan engines over the fuselage, minimizing the front area, the possible momentum caused by one inoperative engine, and therefore the tail size, and reducing the bending moment over the wing when the aircraft is not flying. XI. Structures The structural design of an aircraft is a fundamental aspect of ensuring safety, performance, and efficiency throughout its operational life. The focus of this chapter is on the selection of materials and structural configurations for each primary component: the wing, empennage, and fuselage. A. VN diagram To initiate the structural analysis, the V-N diagram is used to establish the operational boundaries of the aircraft, considering both maneuvering and gust loads—up to 66 ft/s at sea level. This diagram, developed following FAR 23 regulations, helps define the critical speeds and corresponding load factors essential for structural design. Based on the aircraft’s category and weight, the calculated limit load factor is 2.97g. However, accounting for additional loads such as impacts and hydrodynamic forces, a safety factor of 1.2 was applied, resulting in a design limit load of 3.58g. This value remains within the FAA’s maximum recommended limit of 3.8g. The negative load factor is 10 set at -1.0, in line with regulatory guidelines. Further safety margins include a proof load factor of 4.48g (1.25 × limit) and an ultimate load factor of 5.37g (1.5 × limit), which is used for structural design to ensure integrity under both normal and extreme flight conditions. From the V-N diagram in figure 8, several critical speeds are determined: the stall speed (Vs) in clean configuration is 92.51 KEAS, marking the onset of aerodynamic stall at 1g; the maneuvering speed (VA) is 175.04 KEAS, representing the maximum speed for full control input without structural damage; the cruise speed (VC) is 250 KEAS, defining the typical sustained flight speed; the never-exceed speed (VNE) is 337.5 KEAS, acting as the absolute operational limit; and the dive speed (VD) is 375 KEAS, which is the maximum speed before structural integrity is critically endangered. Fig. 8 Survey results for existing aircraft The V-n diagram was also generated as seen in figure 9 for an altitude of 10,000 ft, and several key observations can be made. The stall speed is lower at sea level than at 10,000 ft because less speed is required to sustain the same weight due to the higher air density. The same effect is observed in maneuvering speed, which is a critical parameter defining the lower safe speed limit for performing full control deflections, such as abrupt turns, without risking aerodynamic stall. In summary, all speeds defining the flight envelope tend to be higher at 10,000 ft to compensate for the reduced air density. This advantage allows for significantly higher velocities without the immediate risk of reaching structural limits as quickly as at lower altitudes, even though the aircraft’s structural limits remain unchanged. In summary, the stall speed is 107.65 KEAS, the maneuvering speed is 203.69 KEAS, the never-exceed speed is 337.5 KEAS, and finally the 2 dive speed is 375.00 KEAS. 11 Fig. 9 Survey results for existing aircraft B. Materials Material selection for the fuselage skin considered several aerospace-grade options based on strength-to-weight ratio, corrosion resistance, manufacturability, and cost. After evaluating alternatives such as stainless steel, carbon fiber, fiberglass, titanium, and other aluminum alloys, Aluminum 2024-T3 was selected for its excellent fatigue resistance, high strength-to-weight ratio, and widespread aerospace application as seen in Table 17, where results were consigned. Its availability in the market and familiarity among technicians make it an ideal choice for structural components such as skins and longerons, considering also the harsh operating environment. Al 2024-T3 0.29 0.4 0.3 0.74 0.3 2.03 SS 304 0.34 0.75 0.15 0.08 0.45 1.77 C-Fiber Fiberglass Al 6061-T6 Ti 6Al-4V 0.4 0.26 0.2 0.6 0.33 0.35 0.75 0.25 0.45 0.45 0.15 0.45 0.25 0.09 0.08 0.13 0.15 0.15 0.45 0.45 1.57 1.30 1.63 1.88 Table 17 Selection Matrix results Al–Li 0.35 0.4 0.45 0.11 0.3 1.60 Al 7075-T6 0.38 0.41 0.3 0.09 0.3 1.47 Nevertheless, for the wing structure, each material was selected based on the specific mechanical and operational demands of its corresponding component. The rivets are made of Aluminum 7075 due to its exceptional strength and shear resistance, essential for holding critical structural joints under dynamic loads. Ribs, which shape the airfoil and distribute aerodynamic loads, are made from Aluminum 6061-T6, chosen for its moderate strength, excellent corrosion resistance, and ease of machining. The wing skin uses Aluminum 2024-T3, to match the fuselage skin. For the spar, which is the main load-bearing element of the wing, Aluminum 2024-T4 is used, providing the necessary tensile strength and durability to resist high bending and shear loads while maintaining good toughness. All these decisions are seen in table 18. 12 Materials Fuselage Frames Al 2024-T3 Stringers Al 2024-T3 Longerons Al 2024-T3 Skin Al 2024-T3 Radome Fiberglass Floats Al 2024-T3 Wing Ribs Al 6061-T6 Beam Al 2024-T4 Spars Al 6061-T6 Skin Al 2024-T3 Rivets 7050-T73 Table 18 Materials selection C. Load distribution For the load calculations, the aircraft’s structural weight distribution is defined by the maximum payload of 5,000 lbs, as specified by the Request for Proposal (RFP), and the structural weights of the wing, fuselage, and empennage. These weights are determined based on the materials used and design parameters. Using Corke’s methodology and a fuselage length of 52.33 ft, the analysis defines the magnitude and location of all applied forces, as well as the resulting moments relative to the center of lift, which is located at x/L = 0.553. Moments are measured with respect to this point, with positive values indicating a nose-up tendency. In the aft-tail configuration, the tail lift force provides a negative moment to restore balance and maintain longitudinal stability. Fig. 10 Shear Force and Bending Moment In Figure 10, it can be observed that the shear force remains close to zero for most of the structure, with minor fluctuations, indicating a well-distributed load. In contrast, the bending moment reaches significant values in the central region, attaining a minimum around x/L0.5, which suggests this is the point of maximum bending stress. The fact that the maximum shear and bending moment values occur near the fuselage center is consistent with the lift center location and the overall load distribution, as expected. This implies that the highest internal stresses are concentrated in this area, meaning that reinforcements are required in this region to ensure sufficient strength and stiffness. The maximum shear 13 force is 17,458.87 pounds, while the maximum bending moment reaches-171,000 ft-lbs. These values will be used in the design of the fuselage skin, longerons, and bulkheads, which are the primary structural components of the fuselage. For wing structure C shaped ribs of 3.32 ft long at the wing tip with shank rivets and a C shaped beam were meticulously designed to endure the multiple aerodynamic loadings of the operation such as bending moments and normal forces generated by the lift. In such configuration, 34 ribs of aluminum 6061 T6, with 53 rivets per side will provide enough strength so ensure safety operation, the beams dimension are presented below: Fig. 11 Beam Cross Section D. Structural members The fuselage features a semi-monocoque structure designed to balance strength and weight. In this configuration, the skin plays a primary structural role, mainly resisting tensile stress. To determine the required skin thickness, the fuselage cross-section was approximated as an elliptical shape, taking into account the deadrise and squared characteristics of its geometry. Considering the ultimate strength, the cross-section area dimensions, the maximum bending moment and the ultimate load factor, the minimum required thickness is calculated to be 0.055 in, however commercially available 2024-T3 Alclad Aluminum sheets are easier to find with a 0.056in. Also, the structural elements responsible for handling compressive loads, primarily the longerons and stringers, play a crucial role in preventing skin buckling and distributing longitudinal forces along the fuselage. The longerons are made from Aluminum 2024-T3 to ensure material compatibility with the skin and avoid galvanic corrosion. These components are arranged evenly around the fuselage, spaced every 11.25 degrees, and use a thin-walled circular tube design with a 1.5-inch diameter and 0.065-inch wall thickness. Based on mechanical properties and expected load conditions, the distance between bulkheads was determined to be approximately 13.5 inches, resulting in an estimated 48 frames distributed along the fuselage. Using Ansys, FEA was employed to evaluate, under the most critical conditions, the structural behavior in the most critical areas of the fuselage. Simulations were carried out to represent different frames of the fuselage, using the maximum spacing between them, during cruise, with 1.3 G, a pull-up maneuver at 5.37G and a hard landing at 2G. The results in table 19 show that the deformations generated are minimal compared to the material’s strain limit. The figure 12 displays structural behavior under the 2G load, demonstrating that the frames maintain their stiffness without compromising structural safety. 14 Maneuver Maximum deformation (in) Strain (psi) Cruise 7.87 × 10−5 49.9 Pull Up 0.01575 4379 Landing 0.00059 698 Table 19 Results from the simulation Fig. 12 Structural behavior of the frames Based on these results, it is possible to validate that the rest of the structure can safely withstand the calculated loads. Consequently, the frame spacing is adjusted, widened in regions where the bending moment is near zero, and reduced in key areas where structural demands are greater. Specifically, frames are placed closer together near the center of lift, where shear and bending stress are highest, as well as around door openings and float attachment points. In these critical zones, additional structural elements, such as curved door headers, are integrated to ensure load continuity and preserve the overall structural integrity under demanding operating conditions. On the other hand, the structural design of the wing ensures that it can withstand the highest loads encountered in flight, particularly under the most critical conditions identified from the V-n diagram. These include high-g maneuvers that generate maximum lift, which the wing must sustain not only to support the aircraft but also to compensate for the negative lift produced by the tail and the effects of the fuselage. To handle this distributed load, the wing is modeled as a cantilever beam, which allows for the sizing of essential structural elements such as the main beam, ribs, rivets, and skin. The skin, shaped according to the airfoil, experiences strain due to changing dynamic pressure and transmits these forces through the ribs and rivets to the main beam. The rivets play a critical role in this load transfer, and their selection directly influences the number of ribs required. Based on the strength of the selected aerospace-grade rivets (NAS1097KE5-6), each rib can withstand approximately 19.2 kN. Consequently, 34 ribs are needed to safely distribute the total lift. C-shaped ribs are used to follow the contour of the airfoil and provide structural support to the wing skin, helping distribute aerodynamic loads efficiently to the main spar. The skin thickness is chosen to match the minimum required by the rivet specifications, ensuring adequate strength, choosing AL 2024-T3 sheets with 0.04 in thickness. Additionally, the rib design accounts for both shear and tensile loads, with varying thicknesses at different sections to meet structural demands. A similar structural analysis was conducted for the tail assembly, following the same design philosophy as the wing. The structure is built using also 0.040 in thick aluminum 2024-T3 sheets, and 18 C-shaped ribs are required to ensure adequate load transfer from the skin to the internal structure and to maintain the aerodynamic profile. These 15 ribs are evenly distributed along the span of the tail surfaces, with denser spacing near the root to handle higher stress concentrations. XII. Aerodynamics A. Initial Wing and Empennage Design Given the initial sizing reference of a suggested geometry for the wing and tail design, LLT (Lifting Line Theory) was implemented to design a preliminary wing and then iterate with AVL to obtain the measurements. In order to maximize the aerodynamic efficiency of the aircraft to minimize the thrust required, and therefore the weight and fuel consumption of the engines, a V tail was proposed to ensure stability. The same design process of the wing was used for tail design, the aerodynamic performance with the selected geometry, and measurements are presented as follows: Geometric Data Wing b [ π π‘] 85.18 2 S [ππ‘ ] 604.63 πΆπ πππ‘ [ π π‘] 8.05 πΆπ‘π π [ π π‘] 6.04 MAC [ π π‘] 7.10 AR 12.00 π 0.75 Sweep LE [πππ] 2.00 ππππ [ π π‘] 20.28 AoI [πππ] 1.50 Max. Thickness [ π π‘] 0.97 Min. Thickness [ π π‘] 0.72 Table 20 Geometric Data Tail b [ π π‘] 27.66 2 S [ππ‘ ] 106.67 AR 7.18 πΆπ πππ‘ [ π π‘] 4.26 πΆπ‘π π [ π π‘] 3.41 MAC [ π π‘] 3.85 π 0.80 Sweep LE [πππ] 3.50 AoI [πππ] -1.10 Max. Thickness [ π π‘] 0.38 Min. Thickness [ π π‘] 0.31 Dihedral [πππ] 45 Table 21 Aerodynamic Wing Data For Cruise Airfoil NACA 23012 πΆπ 0.194 πΆπ /πΆπ 30.316 πΆπ 0.007 πΆππ 0.0022 πΆπ -0.084 Lift wing [Lb] 19923.21914 Drag wing [Lb] 661.4525503 Moment [Lb ft] -1003.365933 Table 22 Aerodynamic Tail Data For Cruise Airfoil NACA 0009 πΆπ -0.079 πΆπ /πΆπ -12 πΆπ 0.0066 πΆππ 0.0022 πΆπ 0.035 Lift wing [Lb] -1096.24282 Drag wing [Lb] 84.26873597 Moment [Lb ft] 227.0550333 Table 23 16 B. Control Surfaces Design Fig. 13 Fig. 14 Aileron Design Half Wing Top view 17 Fig. 15 Ruddervator Design Fig. 16 Half tail top view 18 C. Aerodynamic Loadings Fig. 17 Lift Distribution Fig. 18 Shear Force Fig. 20 Flap Forces 19 Fig. 19 Bending Moment Fig. 21 Aileron Forces Fig. 22 Flap Forces 20 Fig. 23 Fig. 24 Flap Forces Ruddervator Forces 21 Fig. 25 Ruddervator Forces D. Performance Fig. 26 Power Vs Speed at each condition 22 Fig. 27 Range Plot Fig. 28 Table 24 Performance Specifications Specification Max Cruise Speed Max Range Takeoff Distance (ISA +18) Landing Distance (ISA +18) Maximum Operating Altitude Maximum Climb Rate Time to Climb (10,000 ft) Maximum Unpowered Range Maximum Endurance @ 10,000 ft Table 25 Fuel Type Trip Fuel Contingency Fuel (15%) Alternate Airport Fuel Final Reserve Fuel Taxi Fuel (1 engine) Takeoff Fuel (T/OFF) Endurance Plot Value 423.78 ft/s 2704 mi 2036 ft 1578 ft 33,500 ft 1570 fpm 7.3 min 35 mi 2 h 51 min Fuel Requirements for a 2h52min (172 min) Flight Estimated Time [min] 172 25.8 25 30 20 – Fuel Weight [lb] 3164.80 474.72 460.00 552.00 184.00 4835.52 Fuel Volume [gal] 472.30 70.90 68.66 82.39 27.46 721.71 XIII. Buoyancy To begin with, the design process focused on estimating the buoyancy parameters for the main float. The first step involved defining the geometry of the master beam section, which refers to the main cross-sectional shape of the float. This geometry is crucial for calculating buoyancy characteristics. According to FAR 23.751 (a), the float must provide at least 80% additional buoyancy above the aircraft’s weight. Using this rule, the displaced (submerged) volume of the main float was calculated for both empty and takeoff conditions. For the empty weight, the displaced volume was estimated at 314 cubic feet (ft³), resulting in a buoyant force of approximately 19,600 lbf. At takeoff weight, the displaced volume increased to about 481 ft³, providing a buoyant force of 30,070 lbf. These values ensure the float can support the aircraft throughout its operational range. 23 Next, the shape of the float’s master section was further defined by calculating the dead-rise angle—the upward slope from the bottom centerline of the float to the sides—and the dead-rise height. The dead-rise angle was found to be 31.1°, and the vertical distance (height) from the keel to the chine was 2.0 feet. Once the main cross-section was established, the next step was to define the stern mirror or transom, which is the vertical or slightly sloped flat surface at the rear of the float. This section significantly affects how water detaches from the float. The stern height was estimated at 5.22 ft, with a height-to-beam ratio of 1.22, and a dead-rise height of 1.21 ft at that point. Using the overall dimensions and assuming a triangular nose shape to simplify calculations, the base (wet) area of the float was found to be 292 square feet (ft²). This area was used to estimate the waterline height, which is important for calculating flotation and stability. For the empty configuration, the waterline height was 1.07 ft, while for takeoff, it was 1.65 ft. The next step involved analyzing hydrodynamic transverse stability, which relates to the float’s ability to resist rolling motions. The center of gravity (CG) of the float’s cross-section was located at 3.99 ft from a reference baseline. The center of buoyancy (CB)—which represents the centroid of the submerged volume—was found to be at 0.72 ft for the empty condition and 1.1 ft for takeoff. With these values, we estimated the metacenter height (GM), a critical measure of stability. The larger the metacenter height, the more stable the float. For the empty case, GM was found to be 1.35 ft, and for takeoff, it was 2.58ft. Figure 29 shows a graphic description for these parameters. Fig. 29 Structural behavior of the frames From this analysis, it became evident that auxiliary floats (outriggers) were needed to improve stability, especially by increasing the metacenter height. According to Gudmundsson’s method, each auxiliary float should displace around 30% of the aircraft’s takeoff weight—not for buoyancy, but to enhance lateral stability. This translated to a required submerged volume of 40.3 ft³ per auxiliary float. To ensure stability in the most critical condition (empty), the minimum beam (width) of each auxiliary float was calculated. A width of 2.69 ft was proposed, considering constraints such as wheel diameter for float storage, and it met the stability criteria. The length of each float was determined using a typical length-to-width ratio of 7:1, resulting in a length of 18.83 ft. The height of the auxiliary floats was based on the NACA 0012 airfoil profile, where the maximum thickness is 12% of the chord (length). This gave a height of 2.27 ft. Using CAD software, the cross-sectional area was estimated at 28.97 ft², giving a total volume of 78.25 ft³ per float—sufficient to meet the stability requirements. Finally, it was decided to align the bottom of the auxiliary floats with the empty weight waterline of the main float. This configuration increases the water contact area during the lightest loading condition and contributes to greater metacenter height and improved stability. The next step involves recalculating overall buoyancy and stability with the auxiliary floats included. Figure 30 shows the auxiliary floats implementation. 24 Fig. 30 Structural behavior of the frames The implementation of auxiliary floats has a significant impact on metacenter height, resulting in a considerable increase and, consequently, improved overall stability. The next step is to determine the longitudinal position of the auxiliary floats relative to the fuselage and the main hull. This requires consideration of both the aircraft’s center of gravity (CG) and the step location to estimate the most appropriate float placement. A suitable approach is to align the center of buoyancy of the auxiliary floats (assumed to coincide with their center of gravity) with the longitudinal center of gravity of the fuselage. To locate the center of gravity of the auxiliary floats, the longitudinal centroid must be calculated. Assuming a rectangular planform (neglecting depth for simplification), the centroid is located at the geometric center in the longitudinal direction. Due to the complex geometry of the transverse section, modeled after an airfoil profile, the centroid position was obtained using CAD software. XIV. Aircraft systems A. Propulsion System Engine type Manufacturer Designation BPR Airflow [Kg/s] Length [mm] Diameter [mm] How many engines are required Mass [kg] Turbofan Pratt & Whitney PW615F * * 1252 556 2 154.4 Thrust Rating Max Continuous at sea level 6495 Takeoff at sea level 6495 Engine Speed Limitations, RPM Max Steady state low rotor (N1] 21830 Max Steady state high rotor (N2] 44040 Transient (20sec) low rotor (N1] 22048 Transient (20sec) high rotor (N2] 44921 FUEL USED JET A-1 Table 26 Air Bleed Max. External bleed available External bleed available 43k ft 17 lb/min 13,4 lb/min Chosen engine characteristics B. Anti-ice & De-ice Given the requirements of an aircraft capable of operating in known icing conditions (RFP), the selection and implementation of an Anti-Ice and De-Ice systems are mandatory. Although one may think that the declared cruise altitude for the aircraft is not high enough for ice formation, it is worth noting that there are regions with constant cold weather climate with sub zero temperatures, most of them remote, in which this aircraft could operate and provide a much needed service as intended by the previously mentioned RFP. 25 1. De-Ice The pneumatic de-ice boots are a proven, mechanically simple solution for keeping the wing and stabilizer leading edges free of accreted ice. Installed directly along the contour of the airfoil, each boot is a flexible rubber membrane that lies flush against the metal skin when deflated, preserving the clean aerodynamic profile essential for lift and control. When ice begins to form—and the pilot or automated ice-detector triggers the system—high-pressure bleed air rapidly inflates the boots, causing them to bulge outward. This sudden expansion breaks the bond between ice and metal, and the airflow immediately strips the fragments away, restoring smooth, ice-free surfaces in seconds. Bleed air for the boots is tapped from one of the engine’s compressor stages and routed through stainless-steel ducts to a Fluiditec MSV98322TB selector valve. Designed to operate reliably from –10 °C up to +60 °C and at pressures up to 10 bar, this valve alternately applies and vents air to each boot section in a timed sequence. After inflation, a SMR2H30-3 relief valve by SMR Technologies opens to exhaust the residual pressure, allowing the rubber to return to its nominal shape and prepare for the next cycle. This combination of engine-driven pressure, precision valve timing, and automatic relief ensures repeated, rapid ice removal without pilot intervention or added aerodynamic penalty. 2. Anti-ice The electric anti-ice system protects windshield and air-data probes by routing thin, heating cables through the probe housings and laminated layers of the windshield. When the ice sensor detects accretion, it signals the electrical control unit—a compact assembly of micro-controllers, relays, and fuses—to energize the heater circuits. Because the wiring is embedded in the critical surfaces themselves, no bulky fairings or airflow penalties are incurred, and the low-mass installation integrates neatly into existing probe and window assemblies. At the heart of each heater element is a silicon-based resistive strip bonded to a thin aluminum backing. Supplied by Rubi Resistances, these custom-width strips (minimum 0.9 mm thickness) operate on 12–24 V DC and withstand temperatures up to 260 °C, with insulation resistance over 5 Mπ to guard against leakage. The Freezing-Rain 0871LH1 sensor by Goodrich (via Campbell Scientific) continuously monitors for ice and automatically toggles the control unit. Pilots also have cockpit switches and annunciators to manually override or verify system status, ensuring clear indications and direct control over anti-ice activation. 3. Hardware characteristics The previously mentioned hardware for both the anti-ice and de-ice systems have te following characteristics: Current (A) Dimension Width (mm) Dimension Height (mm) Dimension Length (mm) Weight (lb) 28 1.5 70.3 38 70.3 28 1.79 25 8 70 Leading edge of wing x x 400 100 200 50.2 Pneumatic Boot 2 Leading edge of stabilizer x x 300 100 140 17.4 Selector Valve Engine bleed-air line Air line before entering pneumatic boot 28 0.2 50 60 120 0.26 x x 50 30 100 1.1 Component Location Voltage (V) Pitot tube and windshield edge Pitot tube and windshield edge Pneumatic Boot 1 Freezing Rain Sensor 0871LH Heater Relief Valve Table 27 Manufacturer System Type 0.0174 Goodrich Electric 2.2 Rebrestisencicas Collins Aerospace Collins Aerospace Fluidtecnic SMR Technologies Electric Pneumatic Pneumatic Electric Pneumatic Anti-Ice and De-Ice System C. Exterior and interior lights The lighting system consists of LED lights, selected for their efficiency, adaptability, and durability, particularly in marine environments. Their sealed design, low power consumption, and high visibility ensure compliance with operational requirements in both airborne and maritime conditions. Additionally, LED lights offer a significantly longer lifespan—approximately 25,000 hours—compared to halogen lights, which typically last around 1,000 hours. The system is connected to the aircraft’s main electrical bus and includes over voltage protection. Emergency lighting is backed by a dedicated battery, and digital dimming allows brightness regulation directly from the cockpit. Automatic sensors activate the landing lights when the landing gear is deployed. All circuits are encapsulated to prevent moisture and salinity intrusion. 26 1. External lights • Navigation Lights: 28 V DC LEDs: red on left wingtip, green on right, white on empennage; visibility greater or equal to 2 NM (ICAO Annex 6). • Anti-Collision Strobes: High-intensity white flashes at 40–100 flashes/min, 30°–60° beam; wingtip and floatmounted; active engine-on to shutdown (14 CFR 91.209). • Landing Lights: 50 W white LEDs in fuselage-wing fairing underside; long-range beam for runway or water surface illumination on approach and rollout. • Taxi Lights: Wide, short-range white beam on nose gear or lower fuselage; supports ground and water-taxi operations. • Float Lights: Red port and green starboard navigation lights plus white anchor lights; compliant with COLREG and aviation regulations. • Wing Inspection Lights: 1–3 W white/amber LEDs near wings and floats; manual or flap/float-deploy activation; vibration-, moisture-, and EMI-resistant. External lighting diagram: Fig. 31 External lighting system External lighting hardware specifications: Light Type Navigation 1 Navigation 2 Navigation 3 Anti-collision Landing Taxi Float (Left) Float (Right) Wing inspection Quantity 1 1 1 3 2 1 1 1 2 Location Left wingtip Right wingtip Top of the empennage Wing tips Under the wings, near the fuselage Main gear Left float Right float Fuselage, near the wing Table 28 Voltage (V) 28 28 28 28 28 28 28 28 28 Current (A) 0.25 0.25 0.25 2.14 1.8 1.8 0.25 0.25 0.175 Weight (lb) 0.2 0.2 0.2 3.66 1.25 1.25 0.2 0.2 0.09 External lights specifications 27 Dimensions (mm) 60 × 69 × 91 60 × 69 × 91 60 × 69 × 91 100 × 80 × 100 220 × 130 × 220 220 × 130 × 220 60 × 69 × 91 60 × 69 × 91 38.1 × 27.9 × 55.8 Color Red Green White White White White Red Green White Manufacturer Fly Wat Fly Wat Fly Wat Honeywell Honeywell Honeywell Fly Wat Fly Wat Aeroleids 2. Internal lights • Cockpit Lighting: Dimmable red/amber LED backlighting for instruments (FAA §25.1385); directional reading lights; emergency LEDs with 90 min battery backup (TSO-C124); IP 67-sealed, corrosion-resistant. • Passenger Cabin Lighting: Adjustable warm/cool white ceiling and sidewall LED tubes; individual reading lights, crew-call panels, air-vent illumination, and safety-sign backlighting; emergency photoluminescent cabin path markers and life-vest lights with 90 min autonomy (FAA §25.1351). • Exit Lights: Adjustable warm/cool white ceiling and sidewall LED tubes; individual reading lights, crew-call panels, air-vent illumination, and safety-sign backlighting; emergency photoluminescent cabin path markers and life-vest lights with 90 min autonomy (FAA §25.1351). Internal lighting diagram: Fig. 32 Internal lighting system Photoluminescence lighting diagram: 28 Fig. 33 Photoluminescence lighting system Internal lighting hardware specifications: LIGHT TYPE QUANTITY Cockpit 1 kit Passenger Cabin 8 LOCATION Reading, instruments, utility, worktable etc. Upper part of the passenger cabin Table 29 VOLTAGE (V) CURRENT (A) WEIGHT (lb) COLOR MANUFACTURER 28 0.18 1.5 White SPTPANEL 28 0.71 5.5 Cyan/White STGAerospace Internal lighting hardware specifications D. Air conditioning In order to regulate humidity, temperature and airflow inside the cabin of the aircraft an air conditioning system must be design to address such need to ensure secure ambient conditions during flight. Its main function will be to bled air from the engines, make it pass through an inter-cooler, and then distribute the air along the aircraft, The presence of an air conditioning system contributes to reduce the thermal fatigue experienced by the crew and improves the air quality inside the aircraft by reducing carbon dioxide and other possible pollutants. The air conditioning system consists of: • Air Cycle Machine (ACM): Hamilton Sundstrand 782790-15 (19.15 lb; 27.9 × 55.8 × 35.6 cm); bleed air → 1st intercooler → compression → 2nd intercooler → turbine expansion → cooled air (+ water separator); turbine-expander core of closed Brayton cooling cycle. • Intercooler: Sacome I-TFM-I integrated in the ACM ( 30 cm long); compact heat exchanger using external airflow to remove compressor heat before turbine expansion. • Water Separator: Honeywell Aerospace 175050; centrifugal deflector design spins airflow to separate and drain moisture, preventing cabin condensation and corrosion. Together, these components ensure that bled engine air is efficiently cooled, dehumidified, and delivered at safe, comfortable conditions throughout all phases of flight. Air conditioning aircraft diagram: 29 Fig. 34 External lighting system Air conditioning hardware specifications: Component Location ACM Air Cycle Machine After the intercooler At the air bleed intake After the ACM Intercooler Water Separator Dimension Width (mm) Dimension Height (mm) Dimension Length (mm) Weight (lb) Manufacturer 355.6 279 558 19.15 Hamilton Sundstrand 150 100 300 4 Sacome 100 100 150 2 Honeywell Aerospace Table 30 Air conditioning hardware specifications E. Inflatable slides The inflatable evacuation slides on this amphibious aircraft serve a dual role: rapid egress in an airborne emergency and life-raft functionality once afloat. Before each flight, the cabin crew arms the system by linking the slide’s inflation hardware to its exit door; in the event of an evacuation, simply opening the door while armed triggers automatic deployment. Within ten seconds, a stored CO2 or N2 gas bottle pressurizes a network of tubes, valves, and chambers to inflate the slide to full form. Each slide incorporates a compact survival kit—flares for signaling and a deployable sun canopy—to sustain occupants in the water. Although FAA 14 CFR §25.809 requires slides for sill heights above 1.8 m, our design’s 1.7 m height makes them optional; nonetheless, their inclusion markedly speeds evacuation and provides built-in life-raft capability, provided at least one trained cabin crew member is on board to arm and disarm the system. In addition to the slide-rafts, the aircraft carries two dedicated marine life rafts to satisfy FAA 14 CFR §25.1411 for aircraft with more than nine seats. These Wicker rafts are stowed near exits for swift deployment and weigh approximately 110 lb each. Upon manual activation of their gas-release valves, they self-inflate when they contact water, offering buoyant, reinforced-walled shelters rated for 10–12 occupants (Fig. 24). Together, the integrated slide-rafts and standalone life rafts ensure compliant, reliable evacuation options for both airborne and waterborne emergencies. 30 F. Public announcement system This system is commonly used in aircraft to clearly distribute announcements or messages made by the crew, whether from the cockpit or the cabin, being highly important, as it delivers essential information to passengers, such as safety instructions, flight updates, and other relevant messages. PA systems generally include audio amplifiers, speakers, and microphones, all of which can be controlled from the cockpit. For this aircraft the selected system is the PA-100 Page/Chime Amplifier from Alto Aviation which is designed for small and medium-sized aircraft. The PA-100 includes: • 3 microphone inputs. • 2 stereo inputs. • 6 command inputs for alarms, cabin alerts, etc. • 4 independent channel outputs, supporting up to 8 integrated speakers throughout the aircraft cabin. The entire system operates on 28V DC, with a maximum current draw of 4 amperes and a weight of 2.3 pounds. It is certified in accordance with FAA TSO C139. The microphone and system controls are in the cockpit, while the speakers are installed in the passenger cabin, typically along the upper fuselage. Fig. 35 Public announcement system G. Avionics 1. Overview and regulation The avionics suite has been centered around Garmin’s integrated ecosystem to support truly single-pilot operations under both VFR and IFR conditions. By standardizing on Garmin hardware and software—spanning the primary flight display (PFD), multifunction display (MFD), and dedicated subsystem gauges—the pilot benefits from seamless data sharing, familiar menus, and consolidated training resources. This approach also simplifies maintenance in remote coastal and island environments, where specialized avionics support may be sparse. All required FAR Part 23, Subparts F and G instruments are included: flight, navigation, and power-plant indicators (§ 23.2615), icing-condition alerts (§ 23.2540), lighting controls (§ 23.2530), and the full suite of VFR/IFR gear (§ 23.2500, § 23.2530). The automated flight-management features and synthetic-vision overlays further reduce workload during complex flight phases—especially critical when transitioning between runway and water operations. Because the amphibious design also functions as a vessel, we’ve added a coast-guard–compliant maritime horn capable of four- to six-second bursts audible up to half a mile, satisfying marine sound-signal regulations for craft over 12 m in length. Rather than compromising the cockpit structure with transparent panels or water-borne viewports, we rely on Garmin’s Synthetic Vision Technology (SVT) built into all NXi series installations. SVT delivers a high-fidelity, 31 three-dimensional representation of terrain, runways, and nearby obstacles—even in reduced visibility—ensuring the single pilot maintains perfect situational awareness during both aerial and surface operations. 2. Component list Component Garmin G1000 NXi Integrated Flight Deck Estimated Unit Price (USD) Garmin GFC 700 Autopilot ∼$20,000 Garmin GIA 64W Integrated Avionics Unit Garmin TAWS Class A Garmin GTS 8000 TCAS II Garmin GTX 335R Transponder with ADS-B Out L3 Harris ESI-500 Standby Instrument ∼$10,000 $52,181.25 Purpose Fully integrated glass cockpit with redundant real-time flight and system data. Digital three-axis autopilot with VNAV and underspeed protection. Combined VHF comm/nav transceiver and WAAS GPS receiver. Terrain awareness/warning with color-coded alerts. Traffic collision avoidance using Mode S and ADS-B. $3,525 Dual-channel transponder with ADS-B Out and diagnostics. $8,089 Independent attitude, airspeed, altitude, heading backup. $59,990 $7,500 Garmin GI 260 AoA Indicator $1,495 Garmin GWX 75 Weather Radar Garmin GMA 1360 Digital Audio Panel Artex ELT 1000 Emergency Locator Transmitter Honeywell AR 120 Flight Data Recorder Garmin GDL 69A SXM Satellite Receiver Garmin GSR 56 Iridium Satellite Data Transceiver Comant CI 292-2 VHF Communication Antenna RAMI AV-529 Navigation Antenna Garmin GA 35 WAAS GPS Antenna RAMI AV-74 Transponder Antenna Sensor Systems S65-536 DME Antenna Artex Whip Antenna Grand total $21,995 Real-time angle-of-attack display for stall prevention. indicator kit. High-definition weather radar with turbulence detection. $1,950 Audio management with Bluetooth, 3D audio, XM Radio. $949 Automatic position beacon for SAR operations. ∼$15,000 Records and stores flight parameters for analysis. $4,658 SiriusXM weather and audio data in the cockpit. $9,445.50 Global satellite comms and data via Iridium network. $403 VHF voice communications antenna. $195 $325 $169 Receives VOR, localizer, and glideslope signals. WAAS-enabled GPS signal antenna. Transmits Mode A/C/S and ADS-B signals. $1,195 Distance measuring equipment signal reception. Included with ELT 1000 Kit $219,064.75. [USD] ELT-specific whip antenna for distress signals. Table 31 Basic avionics component list, price and purpose 32 3. Cockpit component layout Fig. 36 Cockpit component layout H. Core electrical system 1. Overview Modern transport-category aircraft commonly generate three standard power levels: 115–120 V alternating current at 400 Hz for legacy AC loads, and 28 V or 14 V direct current for most contemporary avionics and lighting. Depending on what each subsystem needs, the raw output can pass through transformers, rectifiers, or static inverters that step voltage up or down or even convert AC to DC and vice-versa. From the generator the current flows to one or more primary distribution buses, then on to individual loads, each isolated by a fuse or breaker. At the same time the generators keep the aircraft’s rechargeable batteries topped up; while lead-acid and nickel-cadmium chemistry still dominate, lithium-ion packs are steadily gaining acceptance because of their superior energy-to-weight ratio. For the 19-passenger amphibious aircraft, the electrical architecture is deliberately simple and fully DC. Each Pratt 33 & Whitney PW615F turbofan carries an engine-mounted starter-generator that delivers 28 V to its own main bus through a starter contactor and current-sensing shunt. The left and right buses normally run independent of one another, but a cross-tie relay can close automatically if a generator or engine is lost, allowing the operating side to feed both buses and maintain load balance. A single 24 V main battery (charging to the nominal 28 V bus voltage) is connected downstream of the battery contactor; ground starts and maintenance power come from an external 28 V GPU receptacle wired into the same point in the system (but keep in mind that also the aircraft is equipped with sufficient battery power for non GPU reliance for starting sequences). A dedicated essential—or emergency—bus is fed through isolation devices from at least one main bus and directly from the battery, ensuring that critical functions such as primary communications and standby flight instruments remain energized during any single-source failure. In the unlikely event that both engine-driven generators are inoperative, a compact standby battery supplies this emergency bus for a limited period, giving the crew enough time to complete checklists and land safely. Together, the dual-bus layout, cross-tie capability, and layered battery backup provide redundancy and fault isolation without adding the complexity of an auxiliary power unit. 2. Electrical load calculation Category Max Power (W) Recommended Wiring (AWG) Notes 7 196 AWG 16 for combined feed to each display unit (2.5 A each); AWG 18 for each GIA unit (1 A each). Each display ∼2.5 A Each GIA core ∼1.0 A (avg) 28 5 140 Garmin TAWS (Terrain Awareness) 28 0.5 14 Garmin GTS 8000 TCAS II processor 28 1.7 47.6 System / Equipment Nominal Voltage VDC Garmin G1000 NXi Integrated Flight Deck (2 × 10” Displays GDU + 2 × GIA 64W units) – Primary Flight Display (PFD), Multi-Function Display (MFD), Integrated Avionics 28 Garmin GFC 700 Autopilot (digital flight control system with servos for pitch, roll, yaw) Garmin GTX 335R Transponder (Mode S ADS-B Out, remote) Garmin GWX 75 Weather Radar Garmin GMA 1360D Digital Audio Panel Garmin GI 260 AoA Indicator (stand-alone Angle of Attack display) L3 Harris ESI-500 Electronic Standby Instrument (with internal battery) Rosemount Heated Pitot-Static Probes (2 × probes for capt. & F/O) Max Current Draw (A) 28 1 28 28 2.5 70 AWG 18 (radar power feed). 28 1.5 42 AWG 20 (audio panel power). 28 0.2 5.6 28 0.4 11.2 28 10 280 Artex ELT 1000 (Emergency Locator Transmitter) 28 0 0 Honeywell AR-120 Flight Data Recorder (Cockpit Voice/Flight Recorder) 28 0.3 8.4 MARCO (Boat horn for water operations) 28 3 84 Garmin GA 35 GPS Antenna (active L1 WAAS) 5 0.05 0.25 CI 292-2 VHF Comm Antenna (Comant bent whip) AV-529 Antenna (e.g. second VHF or NAV) RAMI AV-74 Blade Antenna S65-536 Antenna (possibly XM datalink) Artex Whip Antenna (121.5/406 MHz ELT) 0 - 0 A (passive) 0 A (passive) 0 A (passive) 0 A (passive) 0 A (from main bus) - Pratt & Whitney PW615F FADEC (x2, 1 per engine) 28 10 280 Navigation position lights (Whelen/WAT 71105, 5 units) Anti-collision strobes (WAT LED) Landing lights (Honeywell PAR 46 LED) Taxiing lights (PAR 46) Wing-inspection lights (AeroLEDs Starlight) Cockpit FibreLite kit Cabin-passenger LED kit (STG) PA-100 amplifier Anti-ice sensor (Goodrich 0871LH1) 28 28 28 28 28 28 28 28 28 0.25 0.6 2.8 1.4 0.4 1.5 5.5 4 1.5 7 16.8 78.4 39.2 11.2 42 154 112 42 Pitot / windshield heaters 28 7.2 201.6 Avionics Engine Lighting PA System Anti-ice & De - ice AWG 16 to autopilot servo circuits (high-transient current); AWG 20 to AP controller head (low steady current). AWG 22 (fed via G1000 avionics bus – no separate wiring needed if internal). AWG 20 (dedicated circuit from avionics bus; ∼2 A draw). Table 32 AWG 20 (transponder circuit). AWG 22 (very low current device). AWG 22 (short run from emergency/essential bus). AWG14 for each pitot (running constantly regardless of condtion) AWG 22 to ELT control panel (only for arming/trigger, negligible load). AWG 22 (dedicated circuit). 0.5 A (added load) Computing/alerting overhead 0.8–1.0 A during transmition (When a transrepsor consumes more power) 1.5 A (max at high audio output) Nominal ∼0.5–1.0 0.1–0.2 A (nominal) 0.4 A (nominal) ∼0.36 A, 10 W 5 A each ∼0 A during normal ops (has own battery) 0.3 A (nominal) 8 W normal, 10 W AWG 18 (if short run to a 28V horn or PA amplifier). Use dedicated switch/relay. Tiny LNA built into antenna; powered through coax by GIA 64W. Negligible load on 28V bus (∼0.01 A from 28V via internal regulator). Use RG-400 coax; no separate power wire needed. AWG 16 to each engine’s FADEC power 5 A per engine (estimate at full load) (redundant feed from main/emergency bus). Each FADEC channel ∼2–2.5 A; igniters draw Heavy-gauge to handle engine additional ∼3–5 A during start (short duty). transients and dual-channel supply. 3 fuselage + 2 float units (5 × 0.25 A = 1.25 A) Two wing-tip units (0.6 A total) Two retractables (2.8 A total) Two units (0.4 A) Worst-case brightness 4 Elements assume, two pitots, two windshields Electrical load sizing 34 5.0 A (max engaged) Controller ∼0.7 A; 3× servos up to ∼1.5 A Once each component was individually sized according to its datasheet, the total current and power required was obtained, then, a safety factor of 1.5 was applied and the following preliminary current load resulted as: 68.3 1911.25 Amps Watts 102.45 2868.6 Total with 1.5 SF Amps Watts Table 33 Total electric load Total 3. Starter generator selection As mentioned previously, this aircraft is equipped with a starter generator in each engine to provide both a starter source (along a dedicated battery) and a power generation source whilst in flight. Specifically for the PW615F engine there is a vast amount of options for this type of hardware, varying mainly on the amount of current provided as well as certain operational parameters such as altitude, temperature, RPMs and weight. In the case of this aircraft, after an initial load preliminary sizing, the total needed current yields an approximate of 102 amps for continuous operation of avionics, engine, lighting, PA system, de-ice and anti-ice. Therefore, an adequate pick for starter generator for each engine could safely be the MG84 Series with the characteristics shown on the table below. Features MG84 Series Rated Output Continuous 160–200 amps Cooling Self-Cooled Altitude 35,000 ft. Temperature −67 β¦ F to 139 β¦ F Weight 22 lbs. Certification TSO C56b Rated Voltage 30 VDC RPM (Normal Operation) 7,200–12,000 Overload Current 400 amps Over-speed 14,000 Table 34 MG84 Series starter generator key features 35 4. Overall core electrical system diagram Fig. 37 Cockpit component layout I. Aircraft control system The aircraft will retain fully mechanical, mono-yoke, cable-and-pulley primary controls for roll and control rod system for pitch and yaw, assisted only by low-power 28 V DC electric trim motors, Garmin GFC 700 servos and an electromechanical actuator for the flaps. By eliminating hydraulic boosters and fly-by-wire actuators the design stays “non complex”, eases Part-23 certification, cuts extra costs of pumps/lines, and lets, in this case, remote island operators rig or replace a damaged cable with basic hand-tools. 1. Control column A single, center-mounted yoke for the pilot that pivots fore-and-aft for pitch control and rotates left-right for roll control. The yoke’s movements are transferred into cable motions via attached sprockets for the control of the aileron system while the pivoting action movement is transferred to a control rod and moves back and forth a set of control rods for the control of the ruddervator-pitch motion. 36 Fig. 38 Control column single yoke configuration representation for aileron control system 2. Rudder pedals The two rudder pedals serve a dual role: they provide yaw control and, via toe brakes, differential wheel braking on the main gear. Each pedal is mounted on a common shaft under the instrument panel; pressing one pedal forward pulls its control rod aft, while the opposite pedal returns under spring tension to maintain cable and rod tension. The rods run back through fairleads along each fuselage side to the V-tail mixer, where their motion is translated into opposite up/down movement of the ruddervators for yaw. Built into the pedal assembly are master cylinders for the hydraulic brakes: when the pilot toes down on a pedal beyond a small preload, the pedal pushes on its cylinder, applying hydraulic pressure to the corresponding main-wheel brake. This allows tight turns on taxiway and crosswind landing roll control by varying toe pressure. A return spring and adjustable pedal geometry ensure consistent braking feel and that the pedals always center when not in use. Additionally, the pedal shaft features travel stops to limit maximum yaw input and to prevent over-extension of the brake master cylinders. Inspection panels beneath the cockpit allow easy access to pedal bearings, springs, and hydraulic fittings for routine maintenance and rigging checks. 3. Aileron control system The aircraft uses conventional outboard ailerons on each wing to control roll (banking). The pilot’s yoke controls the ailerons via a network of cables, pulleys, and bellcranks that deflect one aileron up while the opposite one goes down, producing the desired rolling moment. The left and right aileron cables run from the yoke area (point A in figure ??) out through the fuselage and wing root to each wing. A set of pulleys (referred as B in figure 40) near the cockpit floor and roof redirects the cables from the vertical plane of the column toward the horizontal plane of the wings. The cables pass through lightening holes, supported by pulley brackets at intervals. Once it reaches the wing’s horizontal plane, both the oncoming left and right cable join in the aileron quadrant (referred as C in figure 40 or as shown in figure 39), here is where the pilot’s input from the control wheel is “translated” into the movement of the ailerons on the wings. 37 Fig. 39 Aileron quadrant reference mechanism The final step, after the moment passes through the quadrant is the actual actuation of the control surface, for this, near each aileron is a bellcrank assembly (B-D assembly near each wings aileron as shown in figure ??) anchored to the wing structure. As the cable from the yoke pulls on one arm of the bellcrank, the bellcrank pivots. The opposite arm of the bellcrank is connected to an aileron pushrod. Thus, the bellcrank converts the cable’s tug into a push/pull on the pushrod. The pushrod then directly attaches to the aileron’s control horn. When the pushrod is pulled, it raises the aileron; when pushed, it lowers the aileron. Fig. 40 Aileron control system logic non-scaled representation 4. Ruddervator pitch control system The aircraft’s V-tail uses two combination surfaces called ruddervators (each surface acts as both rudder and elevator). Achieving proper pitch (elevator) and yaw (rudder) control requires a mechanical mixer that blends the pilot’s yoke and pedal inputs. Each V-tail surface is connected via linkages such that moving the yoke controls them symmetrically (like an elevator for pitch), while pressing the rudder pedals moves them differentially (like a rudder for yaw). When the pilot pushes the yoke forward or pulls it back, the entire control column pivots in the longitudinal (pitch) axis. This motion is transmitted to the ruddervator control system via a set of elevator cables (one cable for nose-up, one for nose-down). In our pulley system, we’ll assume a dual-cable arrangement for pitch: pulling back on the yoke tensions the “up-elevator” cable (and relaxes the “down-elevator” cable), whereas pushing forward tensions the “down-elevator” 38 cable. These elevator cables are routed toward the tail, guided through pulleys along the fuselage. At the aft end of the fuselage (in the tail cone), the elevator cables attach to the ruddervator mixer mechanism. This mixer is typically a bellcrank assembly that can impart equal deflections to both left and right ruddervators simultaneously. For pitch, the mixer moves both ruddervator surfaces in the same direction together. Fig. 41 Fig. 42 Mechanical mixer diagram for elevator control Elevator control system logic non-scaled representation 39 5. Ruddervator yaw control system Yaw inputs from the rudder pedals are transmitted through two parallel control rods—one on each side of the fuselage—that attach to the differential input of the V-tail mixer. Pressing the right pedal pulls the right rod aft and pushes the left rod forward, causing the mixer to deflect the left ruddervator up and the right down, thus yawing the aircraft to the right. The return-spring assembly at the pedals maintains tension and re-centers the controls when feet are lifted. Fig. 43 Fig. 44 Mechanical mixer diagram for rudder control Rudder control system logic non-scaled representation 40 J. Fuel system Given the fuel requirements for maximum endurance operation, the fuel tank must hold at least 721.71 gallons. To ensure operational margin, the tank will be designed with a 10% volume surplus, resulting in a total capacity of 793.88 gallons. Since the chord at the wing tip is known, the thickness and depth of the wing are fixed and will serve as reference dimensions for the tank design. With the volume (3.005 π 3 ), height (0.2 π), and depth (1 π) defined, the remaining variable is the tank width, which is calculated to be 7.515 meters. To meet these requirements, two fuel tanks will be installed within the wing using a wet wing configuration. Each tank will have dimensions of 1 m depth, 20 cm height, and 7.515 m width. Both tanks will be installed in the wing section outside the fuselage. K. Landing gear The landing gear design for the amphibious aircraft was developed using methodologies outlined by Raymer, Gudmundsson, and Mohammad. Various configurations were evaluated, including integration within hydrofoils, nacelles, wings, the passenger cabin, and external mountings, each one presenting trade-offs in aerodynamic performance, structural complexity, weight distribution, and maintenance accessibility. After several iterations, a tricycle configuration was selected, retractable into the fuselage-mounted floats. The mechanism, similar to the A6 Grumman, is a forward-folding, articulated strut configuration. An initial sizing of the gear determined the vertical clearance from the ground to the fuselage as 1.97 ft, with the total gear height approximated at 2.62 ft and the center of gravity (CG) estimated to be 7.58 ft above the ground. The final distance between the CG and the main gear contact point was calculated to be 2.96 ft, ensuring stability against tipback during rotation. Stability analyses confirmed that the aircraft maintains a tipback angle of 14.87°, which exceeds the operational rotation angle plus the required margin, confirming that the aircraft will not tip backward on the ground. Additionally, a clearance angle of 22.31° was computed between the tail and ground at rotation, indicating no risk of tail strikes during takeoff. The wheelbase was determined to be 19.75 ft, while the track—the lateral distance between main wheels—was set at 7.39 ft. This configuration satisfies ground lateral stability criteria, with an overturn angle of 48.41°, comfortably within the accepted range (between 25° and 63°). Steering performance was assessed by analyzing how nose gear angle affects turning radius. A nose gear angle of up to 22° was found to provide sufficient maneuverability for a range of runway conditions, reinforcing the decision to include a nose wheel steering system. The load distribution analysis showed that the main landing gear supports approximately 91% of the aircraft’s static weight (15,250.97 lb), while the nose gear supports 11–15% (1,783.95 to 2,555.24 lb), aligning with standard design guidelines. The gear must also endure dynamic loads up to 39% of total aircraft weight during landing and taxi operations. To verify proper load sharing, ratios between the CG and landing gear positions were computed. Both the forward and aft CG limits fell within acceptable values, confirming that the aircraft’s balance and stability are well managed under static and dynamic conditions. Based on this analysis, the required tire dimensions were established. The main tires are estimated at 26.57 inches in diameter and 7.41 inches wide, while the nose tires are proportionally smaller, at 18.6 inches diameter and 5.19 inches wide. Using the Goodyear Aviation catalog [3], suitable tires were selected as seen in table 35: Position Main Gear Nose Gear Tire Model Rated Static Load 27.7 × 8.75–14.5 21,500 lb 18 × 5.7–8 9,000 lb Table 35 Tire selection Dimensions (in) ∅ 27.7 × 8.75 ∅ 18.6 × 5.19 For absorbing impact energy during landings, oleo-pneumatic shock struts were selected due to their proven effectiveness in dissipating energy on unprepared surfaces. A vertical touchdown velocity of 13 ft/s was used as a design condition, in line with typical STOL aircraft operating on soft or rough runways. The resulting stroke of the shock absorber was calculated at 12.15 inches, above the recommended minimum of 8 inches, and the total oleo length was set at 30.39 inches. The external diameter was computed as 2.9 inches, assuming standard pressure conditions. Finally, braking energy was evaluated based on a landing weight equal to 70% of the aircraft’s maximum takeoff weight and a stall speed of 107.81 ft/s. The resulting braking energy was distributed equally between the two main wheels, yielding 1,059,602 ft·lb per wheel, a value comfortably supported by the selected braking system. Figure references confirmed the selected tires allow proper integration of the brakes without dimensional conflict. 41 XV. Weight and Balance Component Weight and Coordinate Data Component Weight lb X in Wx in.lb Y in Wy in.lb Z in Wz in.lb Wing Empennage Fuselage Nose Landing Gear Main Landing Gear Engine Nacelles Fixed Equipment Fuel System Avionics Hydraulic System Air contidioning and anti ice 674.98 383.64 5622.14 10.41 62.19 1673.07 182.13 1433.40 214.40 265.21 76.05 197.36 211.02 586.25 314.96 59.00 314.26 195.10 183.29 314.96 211.02 39.37 314.96 174.52 142433.84 224909.08 1770750.52 614.29 19542.33 326415.89 33383.05 451464.17 45243.25 10441.16 23952.57 34442.47 0 0 0 0 0 0 0 0 0 0 0 0 0 0 0 0 0 0 0 0 0 0 0 0 118.10 216.13 68.30 11.81 11.81 131.40 131.40 68.30 115.94 68.48 23.62 118.89 79714.89 82916.16 383992.44 122.96 734.41 219841.35 23932.33 1433.40 24857.84 18161.30 1796.29 23463.59 Empty Weight We 10794.98 286 3083592.60 0 0 89 860966.97 TFO Fuel 1 Crew 3 pax 3 pax 3 pax 3 pax 3 pax 2 pax 2 pax Baggage 44.00 1519.87 175.00 525.00 525.00 525.00 525.00 525.00 350.00 350.00 748.00 211.00 211.02 137.79 185.48 213.48 241.48 269.48 297.48 325.48 353.48 393.25 9284.00 320722.97 24113.25 97377.00 112077.00 126777.00 141477.00 156177.00 113918.00 123718.00 294151.00 0 0 0 40 40 40 40 40 20 20 0 0 0 0 20909 20909 20909 20909 20909 6944 6944 0 118.10 115.94 93.89 93.89 93.89 93.89 93.89 93.89 93.89 93.89 68.31 5196.40 176213.73 16430.75 49292.25 49292.25 49292.25 49292.25 49292.25 32861.50 32861.50 51092.14 Take-Off Wht 16606.85 277 4603384.82 7 118433 91 1422084.24 42 XVI. Financial Report Results and Route Analysis A. Cost analysis report results Applying Gudmunson’s methodology, the development and production costs for the aircraft and the man-hour requirements at each stage were estimated. Development costs—covering engineering, tooling, manufacturing, certification, support, and flight testing—represent the upfront investment before entering serial production. We also quantified the total man-hours needed for engineering, tooling, and manufacturing, providing a clear view of the workload and resources demanded by each phase. Production costs per unit incorporate materials, manufacturing, quality control, avionics, engines, and other key components. From these figures, we derived an estimated sale price and calculated the number of aircraft required to reach the break-even point. The detailed results of these analyses follow. 43 1. Development costs DEVELOPMENT COSTS Description Value (USD) Quantity Discount Factor (QDF) 1,0156 Total Development Cost $1 583 720 458.75 Engineering Cost $157 098 849.08 Tooling Cost $57 524 799.29 Manufacturing Cost $122 114 661.74 Development Support Cost $94 509 612.29 Flight Testing Cost $421 669 637.85 Certification Cost $730 802 898.51 Table 36 Development costs Hours – – 1 023 910.87 605 647.88 2 507 076.08 – – – 2. Production costs PRODUCTION COSTS Description Value (USD) Production Cost per Unit $6 226 278.36 Materials Cost $2 094 373.71 Manufacturing Cost $745 961.03 Quality Control Cost $116 369.92 Avionics Cost $219 064.75 Engine Cost $2 782 913.48 Maintenance Cost (Per Year) $29 962.40 Estimated Selling Price $7 471 534.03 Table 37 Production costs 3. Break even analysis BREAK-EVEN ANALYSIS Description Value Units Required to Break Even 99,47 Table 38 Break-even analysis 44 4. Cost summary COST SUMMARY Description Value (USD) Total Fixed Cost $730 802 898.51 Selling Price per Unit $7 340 273.84 Variable Cost per Unit $124 525.57 Table 39 Cost summary B. Operational cost summary OPERATIONAL COST SUMMARY (Value USD) Description Value (USD) Annual Storage Cost (per year) $10 200 Annual Fuel Cost (per year) $304 200 Insurance Cost (per year) $112 573.01 Annual Inspection Cost (per year) $500 Annual Engine Overhaul Fund (per year) $10 400 Annual Loan Payment $884 968.51 Total Annual Cost $1 210 268.51 Cost per Flight Hour $1 163.72 Table 40 Operational cost summary C. Cost composition graph Fig. 45 Development cost breakdown 45 Fig. 46 Cost per unit distribution Fig. 47 Operational costs D. Route analysis The following section details both national and international routes the aircraft could cover without exceeding its operational limits. These routes might pose a business importunity given the fact tat they each have a significant advantage for the designed aircraft. 46 Origin Destination Distance (NM) Time (h) Operation Type Cartagena Providencia 210 0.84 Tourism / Scheduled San Andrés Corn Island 220 0.88 International / Charter International Buenaventura Bahía Solano 130 0.52 Logistics / Tourism Amphibious advantage Leticia Tabatinga 4 0.02 Medical / Logistics Direct competition Capurganá Puerto Obaldía 8 0.03 Social / Logistics Cartagena Nuquí 230 0.92 Tourism / Charter Guapi Tumaco 190 0.76 Social / Logistics San Andrés Bluefields 250 1 International / Charter International Mitú Inírida 180 0.72 Social / Connectivity Direct competition Table 41 Classification Direct competition Amphibious advantage Amphibious advantage Amphibious advantage Potential route analysis XVII. References 47 Justification / Observations Tourist route operated by Twin Otter; potential advantage with direct service to bays. No regular connection; potential opening of an international amphibious market. Difficult access; reduces travel time by bypassing airports. Short route, rapid medical and light-cargo connection between the two. Water access between isolated communities; ideal for local cargo and passenger service. Access to areas without airports; major regional tourist attraction. Coastal towns with limited land or air access. Utilizes full range; connects insular areas of Nicaragua and Colombia. Jungle region with Twin Otter presence; potential to serve social and medical transport.
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