**3. Flight test vehicle structural model development**

Initially, several low-cost, small model aircraft were considered for finite element analysis and simulation, with multifunctional lightweight composite panels replacing part of the wing and fuselage structure. A remotely piloted small aircraft was selected with a 127 inch wingspan and a takeoff weight of 16 lb. Adequate details about the internal structure and fabrication of this model airplane were not known, so a notional FEM of this small aircraft was quickly developed for initial structural analysis with design flight loads. **Figure 5** shows a preliminary structural model development of a similarly sized small hobby model airplane, which offered an initial low-risk candidate for flight testing of the M-SHELLS specimen. A typical wing FEM with a standard two-spar and rib configuration was initially developed. This structural arrangement would enable easy integration of small test coupons, between the two spars in the inboard section, close to the electric motor in the fuselage nose. The test specimen could also be integrated into the fuselage floor.

**Figure 6** shows the wing deflection and strain distribution from initial structural analysis of the wing in level flight. The analysis assumed front and rear spar thicknesses of 0.15 inch with advanced composite material properties [5]. The linear elastic property values used for the front and rear spar are as follows: Young's modulus 9,750,000 psi, shear modulus 2,570,000 psi, and mass density 0.06 lb/in3 . The wing, fuselage, horizontal tail, and vertical tail skin thicknesses were 0.04 inch and were made of standard thermoplastic material. The linear elastic properties are as follows: Young's modulus 290,075 psi, shear modulus 47,250 psi, and mass density 0.04 lb/in3 . The wing deflections and skin strain distributions shown are with a fixed wing root and a 16 lb lift load, distributed elliptically along the wing. The maximum deflection and nodal strain were 1.95 inches at the wing tip and 0.00106 at the wing root, respectively. With this two-spar wing construction, the maximum wing-tip deflection and strain values at level cruise flight were considered high for a model airplane. The two-spar wing FEM weight was calculated to be 4.63 lb. The fuselage weight, with empennage, was calculated to be 3.8 lb.

When NASA Langley acquired two UASUSA-manufactured remotely piloted aircraft named "Tempest" for the planned flight test, additional information on the internal construction of the physical model was available. A *Tempest* model was dismantled to observe the internal construction at the wing root. The weight of each component of the disassembled model was also measured. Since the material properties of the *Tempest* wing and other model parts were not known, a bench test

*Environmental Impact of Aviation and Sustainable Solutions*

The proposed M-SHELLS research goals were to develop test specimens and subcomponents, integrate them into a small test vehicle structure, and conduct low-risk flight tests. The M-SHELLS test coupons in the form of honeycomb panels were fabricated and tested by Russell Smith (LaRC) and Brett Bednarcyk (GRC) for mechanical and electrical properties. **Figure 3** shows the normal compression load shakedown test of a small, stabilized aluminum honeycomb coupon fabricated for mechanical property assessment. The compressive crushing strength and compressive modulus were computed and compared with the published characteristics of a Hexcel 1/4-5052-0.002 honeycomb. The flatwise compression modulus of the aluminum honeycomb coupon with 1/4-inch cell and 0.002-inch foil thickness is 139,000 psi and the crushing strength is 436 psi. The published in-plane shear modulus of the Hexcel 1/4-5052-0.002 honeycomb is 66,000 psi and the shear strength is 300 psi in the length direction. In the width direction, the in-plane shear modulus is 30,000 psi and the shear strength is 120 psi. Since the normal compression strength test result and Hexcel published data were very close, the mechanical properties of Hexcel honeycomb were used by Olson and Ozoroski [2] for the initial structural and multifunctional performance benefit analysis of the N3CC derivative with hybrid-electric propulsion. They also accounted for the additional weight of core material required to complete the energy storage functionality. **Figure 4** shows the in-plane tensile load versus extension plot from an initial tensile test of an early M-SHELLS active coupon prototype with anode/cathode elements and electrolytes. The honeycomb test coupon dimensions were 6.0 inch (150 mm) in length, 2.0 inch (50.8 mm) in width, and 1.0 inch (25.4 mm) in depth. The face-sheets were 0.002 inch thin aluminum foil. The electrical tests were conducted at NASA Glenn Research Center. Considering only the linear part of the deformation, a 90 lb (400 N) load produces an extension of 0.6 mm. Thus, relative to the unloaded specimen, the linear elastic strain was 0.004 at the 90 lb (400 N) load. The specimen yielded beyond the 400 N load and developed a crack at 480 N. The linear Young's modulus (stress/strain) was computed to be 11,188 psi

). The corresponding in-plane shear modulus was 4024 psi for

the Poisson's ratio of 0.39. The in-plane tensile and shear modulus computed from the coupon test results were very low for flight application. Hence, for the present analysis, additional outer face-sheets were added on each side to add strength to the honeycomb core (**Figure 1b**). Several detailed finite element models (FEM) of three flight vehicles were developed having certain fuselage areas replaced with this reinforced composite panel having a honeycomb core. Structural analyses of these models are described. The complete summary of all material properties used in this

*Normal-compression load shakedown test of a small, stabilized aluminum honeycomb coupon fabricated for* 

**2. M-SHELLS coupon test**

**6**

**Figure 3.**

*mechanical property assessment.*

(77.52 × 106 N/m2

**Figure 5.** *Preliminary structural model development of the two-spar wing airplane.*

**Figure 6.** *Wing deflection and strain of the two-spar wing model airplane.*

was performed to evaluate the wing deflection and stiffness under a simulated lift load. Gregory Howland and David Hare performed the bench load-deflection test at the NASA Langley model shop on a layout table. The loading configuration was based on the test setup scheme shown in **Figure 7**. The model was inverted and then leveled and supported by two foam blocks. The wing load application points were positioned at 24 inches from the centerline. Eight-pound weights were placed on the right and left wings symmetrically at those reference points. The average wing-tip displacement was ~0.94 of an inch. The load was removed from each wing and then the loading was repeated. The second time, the average wing-tip deflection was 0.96 of an inch. The inset photos in **Figure 7** show the bench test arrangement in the NASA Langley model shop.

Upon close examination of the model with the canopy removed, it was observed that the *Tempest* wing is constructed as two symmetric pieces of hollow, molded composite that are joined together with a short central stub-spar and two solid root-rib pieces, each 2 inches wide. **Figure 8** shows the *Tempest* wing construction. A new finite element model of the wing was developed to represent this construction. The central stub-spar and two wide ribs were modeled with solid advanced composite material properties as before. The molded fiberglass skin of the two wings was modeled as 0.025 inch thin composite material. The rest of the model used custom thermoplastic material.

The horizontal tail skin and ribs were modeled as 0.02 inch thin molded thermoplastic. The fuselage and vertical tail skins and ribs were modeled with 0.04 inch thin thermoplastic. The horizontal and vertical tail twin-spar thicknesses were

**9**

**Figure 7.**

**Figure 8.**

system, and motors.

*Structural Analysis of Electric Flight Vehicles for Application of Multifunctional Energy Storage…*

0.10 inch and 0.15 inch, respectively. **Figure 9** shows the wing deflection and nodal strain distributions from the FEM analysis with level flight load, assuming a 16 lb takeoff gross weight. See Appendix A for all the material elastic properties and density used in this chapter. With the improved FEM of the wing structure, the wingtip deflection was 1.11 inch and the maximum strain at the wing root was 0.00067. The strain values were noted to be well within the allowable limits. The wing-tip deflection was closer to the experimental results than the preliminary FEM analysis results with the two-spar wing (**Figure 6**). This improved FEM analysis result was

**Table 1** shows the measured component weights of the test vehicle and estimated weight for the initial two-spar wing model and the improved model of the **Tempest** wing. Some of the structural component weights and the electronic system weight inside the fuselage could not be measured separately, since the fuselage and vertical tails are molded as a single part. Hence, the weights of those components are grouped together in **Table 1**. The two-spar wing weight was estimated to be 4.63 lb. With the better FEM of **Tempest**, the estimated total wing weight of 3.54 lb is closer to the measured combined weight of 3.46 lb for its right and left wings and stub spar. The measured fuselage weight, 5.62 lb, included the co-molded vertical tail and electronic components inside the fuselage. It compared well with the improved FEM combined weight of the fuselage and vertical tail, including an estimated 2 lb weight for electronic components, telemetry

considered satisfactory for the structural component weight estimation.

*Wing deflection test of the tempest aircraft with 16 lb total lift load on the wing.*

*Structural model and wing root internal detail of the tempest aircraft.*

*DOI: http://dx.doi.org/10.5772/intechopen.86201*

*Structural Analysis of Electric Flight Vehicles for Application of Multifunctional Energy Storage… DOI: http://dx.doi.org/10.5772/intechopen.86201*

**Figure 7.** *Wing deflection test of the tempest aircraft with 16 lb total lift load on the wing.*

**Figure 8.** *Structural model and wing root internal detail of the tempest aircraft.*

0.10 inch and 0.15 inch, respectively. **Figure 9** shows the wing deflection and nodal strain distributions from the FEM analysis with level flight load, assuming a 16 lb takeoff gross weight. See Appendix A for all the material elastic properties and density used in this chapter. With the improved FEM of the wing structure, the wingtip deflection was 1.11 inch and the maximum strain at the wing root was 0.00067. The strain values were noted to be well within the allowable limits. The wing-tip deflection was closer to the experimental results than the preliminary FEM analysis results with the two-spar wing (**Figure 6**). This improved FEM analysis result was considered satisfactory for the structural component weight estimation.

**Table 1** shows the measured component weights of the test vehicle and estimated weight for the initial two-spar wing model and the improved model of the **Tempest** wing. Some of the structural component weights and the electronic system weight inside the fuselage could not be measured separately, since the fuselage and vertical tails are molded as a single part. Hence, the weights of those components are grouped together in **Table 1**. The two-spar wing weight was estimated to be 4.63 lb. With the better FEM of **Tempest**, the estimated total wing weight of 3.54 lb is closer to the measured combined weight of 3.46 lb for its right and left wings and stub spar. The measured fuselage weight, 5.62 lb, included the co-molded vertical tail and electronic components inside the fuselage. It compared well with the improved FEM combined weight of the fuselage and vertical tail, including an estimated 2 lb weight for electronic components, telemetry system, and motors.

*Environmental Impact of Aviation and Sustainable Solutions*

*Preliminary structural model development of the two-spar wing airplane.*

*Wing deflection and strain of the two-spar wing model airplane.*

was performed to evaluate the wing deflection and stiffness under a simulated lift load. Gregory Howland and David Hare performed the bench load-deflection test at the NASA Langley model shop on a layout table. The loading configuration was based on the test setup scheme shown in **Figure 7**. The model was inverted and then leveled and supported by two foam blocks. The wing load application points were positioned at 24 inches from the centerline. Eight-pound weights were placed on the right and left wings symmetrically at those reference points. The average wing-tip displacement was ~0.94 of an inch. The load was removed from each wing and then the loading was repeated. The second time, the average wing-tip deflection was 0.96 of an inch. The inset photos in **Figure 7** show the bench test arrangement in the

Upon close examination of the model with the canopy removed, it was observed that the *Tempest* wing is constructed as two symmetric pieces of hollow, molded composite that are joined together with a short central stub-spar and two solid root-rib pieces, each 2 inches wide. **Figure 8** shows the *Tempest* wing construction. A new finite element model of the wing was developed to represent this construction. The central stub-spar and two wide ribs were modeled with solid advanced composite material properties as before. The molded fiberglass skin of the two wings was modeled as 0.025 inch thin composite material. The rest of the model

The horizontal tail skin and ribs were modeled as 0.02 inch thin molded thermoplastic. The fuselage and vertical tail skins and ribs were modeled with 0.04 inch thin thermoplastic. The horizontal and vertical tail twin-spar thicknesses were

**8**

NASA Langley model shop.

**Figure 5.**

**Figure 6.**

used custom thermoplastic material.

**Figure 9.**

*Wing deflection and strain of the improved finite-element model of the test vehicle in level flight.*


**Table 1.**

*Comparison of component weights of the tempest test vehicle, initial two-spar wing model, and improved tempest FEM.*

The performance goal for the M-SHELLS development was to demonstrate a specific power of 1000 W/kg at an energy density of 75 Wh/kg. The flight test goal was to augment the existing Li-Po battery with 33% of the required energy for 30 minutes of flight or, equivalently, to supply the full electrical energy for 10 minutes of level flight. The Li-Po battery capacity is 7600 mAh and it provides 7.4 volts with two 3.7 volt cells in series. With a gross weight of 2.3 lb (1.04 kg), the energy density of the Li-Po battery is 55 Wh/kg. The ideal power required by the aircraft at cruise is computed from weight × velocity/(L/D), where L/D is the lift-to-drag ratio. Considering the propeller and motor efficiencies, the total power required to be supplied to the electric motor spinning the propeller is:

Power Required <sup>=</sup> weight <sup>×</sup> velocity/[L/D <sup>×</sup> (propeller efficiency) × (motor efficiency)] (1)

For the *Tempest* test vehicle, let us assume a baseline cruise weight of 20 lb (88 N), a cruise velocity of 40 mph (17.9 m/s), and a typical L/D of 20. Assuming a motor efficiency of 85% and a propeller efficiency of 80%, the power required = 88 × 17.9/ (20 × 0.85 × 0.80) = 116 W and the energy required for 10 minutes of level flight is (116 × 10/60) = 20 Wh. Hence, ideally, 0.58 lb (20/75 kg) of M-SHELLS material could provide full power for 10 minutes of level flight. The actual weight of the M-SHELLS power package would depend on the flight test voltages and current demand of the electric motor and the ability to package each unit in suitable series and parallel configurations to match the available power supply and required power demand.

**11**

738 lb (335 kg).

*Structural Analysis of Electric Flight Vehicles for Application of Multifunctional Energy Storage…*

The structural deflection and nodal strain distribution from the FEM analysis

of mid-fuselage floor

results of the *Tempest* vehicle with a lightweight M-SHELLS composite panel replacing the fuselage floor are shown in **Figure 10**. The five-layer bonded sandwich panel consisted of 0.02 inch thermoplastic sheet for insulation on the outer faces, 0.002 inch aluminum sheet on the inner faces and 1.0 inch deep honeycomb M-SHELLS core. The original fuselage floor weight was 0.32 lb. One stack of this

would weigh 1.25 lb. The mid-fuselage floor composite, multifunctional panel would provide both structural integrity and supply electrical energy to supplement

Under the Scalable Convergent Electric Propulsion Technology Operational Research (SCEPTOR) project, the X-57 Maxwell test vehicle wing is presently being constructed at NASA Armstrong Flight Research Center. **Figure 1e** showed the NASA X-57 Maxwell experimental test aircraft concept [4] with a distributed electric propulsion system featuring 12 electric-motor-driven propellers on an innovative high-lift wing. The X-57 Maxwell vehicle will test the performance of this specially designed wing with distributed electric propulsion in order to evaluate

**Figure 11** shows the weight breakdown of the NASA X-57 Maxwell experimental test aircraft. The original wing of the Italian *Tecnam P2006T* aircraft will be replaced with a specially designed distributed electric propulsion wing with 12 electric-motor-driven propellers. The wing-tip propellers help reduce the induced drag from the tip vortex. The synchronized motors are powered by a 358 kg Nickel-Cobalt-Aluminum (NCA) battery pack. The electric power system is organized into eight battery modules, split into two packs with 4 battery modules and a control module each. Cooling is provided through 18,650 cells spaced evenly, 4 mm apart. The NCA cells provide sufficient energy density and the required discharge rate for the flight test mission. Each pack supplies 47 kWh of useful energy, with a peak discharge power of 132 kW. The total battery package weight is estimated to be 790 lb (358 kg), or 26% of the total aircraft takeoff gross weight of 3006 lb (1364 kg). The aluminum fuselage weight is 302 lb (136 kg), and the total estimated structure weight without the landing gear is

**Figure 12** shows initial power requirement estimates for the standard mission of the X-57 Maxwell [6] flight test vehicle. The energy requirement for each phase of the mission is obtained by integrating the power requirement over time

five-layer sandwich energy storage panel replacing 180 in2

*Tempest FEM analysis with M-SHELLS composite panel fuselage floor.*

the existing Li-Po battery of this vehicle.

**Figure 10.**

**4. NASA X-57 Maxwell test vehicle**

mission benefits for this class of vehicle.

*DOI: http://dx.doi.org/10.5772/intechopen.86201*

*Structural Analysis of Electric Flight Vehicles for Application of Multifunctional Energy Storage… DOI: http://dx.doi.org/10.5772/intechopen.86201*

**Figure 10.** *Tempest FEM analysis with M-SHELLS composite panel fuselage floor.*

The structural deflection and nodal strain distribution from the FEM analysis results of the *Tempest* vehicle with a lightweight M-SHELLS composite panel replacing the fuselage floor are shown in **Figure 10**. The five-layer bonded sandwich panel consisted of 0.02 inch thermoplastic sheet for insulation on the outer faces, 0.002 inch aluminum sheet on the inner faces and 1.0 inch deep honeycomb M-SHELLS core. The original fuselage floor weight was 0.32 lb. One stack of this five-layer sandwich energy storage panel replacing 180 in2 of mid-fuselage floor would weigh 1.25 lb. The mid-fuselage floor composite, multifunctional panel would provide both structural integrity and supply electrical energy to supplement the existing Li-Po battery of this vehicle.
