# Design and Analysis of VTOL Operated Intercity Electrical Vehicle for Urban Air Mobility

^{*}

## Abstract

**:**

## 1. Introduction

## 2. Literature Study

#### Study of Similar Aircrafts

## 3. Layout

#### 3.1. Initial Design and Discussion

#### 3.2. Overview of the Present Design

#### 3.3. Configuration Selection

#### 3.4. Wing Planform

#### 3.5. Fuselage

#### 3.6. Landing Gear

#### 3.7. Aerodynamic Control Surfaces

#### 3.8. Propulsion

- Since the number of propellers in our design is six (four mounted on the wing and two on V-tail), the required thrust is obtained by increasing the diameter of the blades of the propellers. If we opt for larger diameter blades, we have to increase the area of the duct. This increases the wetted surface area which in turn increases the drag. Furthermore, the interaction of duct vortices might reduce the efficiency of the rear propellers. The construction of the duct and its support structure increases manufacturing costs and also increases the total gross weight of the eVTOL aircraft.
- In a ducted fan, the manufacturing tolerances are essential; the tighter it is, the more efficient the fan is. However, as it becomes tighter, the fans are susceptible to hitting the duct due to the flexing of blades. The pitch control and RPM control become difficult to implement due to the ducts.
- In the event of a crosswind on a ducted fan, it results in a sudden pitching moment due to the lip of the duct and the asymmetric inflow to the rotor makes the aircraft lose control and stability [26].

## 4. Mission Profile and Preliminary Sizing Calculations

- Range: 500 km;
- Crew: one pilot (90 kg with baggage);
- Payload capacity: 500 kg;
- Air-taxi mode: Four (04) Passenger and 0.5 m
^{3}cargo hold; - Air-cargo mode: 1.5 m
^{3}cargo hold by removing passenger seats; - Cruise and maximum speed: 200 and 250 km/h TAS;
- Powerplant: green propulsion using non-fossil fuel;
- Operational constraint: take-off and land from building rooftop;
- Enroute constraint: maintain separation of >600 m;
- Entry to market: before 1 January 2031.

#### 4.1. Mission Profile Planning

#### 4.2. Initial Sizing, Methodology and Payload-Range Diagram

#### 4.2.1. Methodology for Sizing

#### 4.2.2. Iteration Process in the Estimation of the Weight of the Wing

_{wing}= 0.036 S

_{w}

^{0.758}W

_{fw}

^{0.0035}q

^{0.006}λ

^{0.04}

_{wing}is the weight of the wing, S

_{w}is the trapezoidal wing area, W

_{fw}is the weight of fuel in the wing, q is dynamic pressure at cruise and λ is the taper ratio.

#### 4.2.3. Iteration Process in the Estimation of the Weight of the Fuselage

_{fuselage}= (0.052 S

_{f}

^{1.086}L

_{t}

^{−0.051}q

^{0.241}) + W

_{press}

_{fuselage}is the weight of the fuselage, S

_{f}is fuselage-wetted area, L

_{t}is tail length (wing quarter MAC to tail quarter MAC) and W

_{press}is weight penalty due to pressurization. On estimation, the weight of the fuselage was found to be impractical for the given dimensions; hence, it was calculated using the simplified version of the same equation. Furthermore, the same was analysed using the OpenVSP tool and the desired result was obtained.

#### 4.2.4. Iteration Process in the Estimation of the Weight of V-Tail and Flight Controls

_{horizontal tail}= 0.016 q

^{0.168}S

_{ht}

^{0.896}λ

_{h}

^{−0.02}

_{flight controls}= 0.053 L

^{1.536}B

_{w}

^{0.371}

_{ht}is a horizontal tail area, λ

_{h}is taper ratio of the horizontal tail, L is fuselage structural length, B

_{w}is wingspan. Equations (3) and (4) were helpful in determining the desired weight of the V-tail and the flight controls.

#### 4.2.5. Iteration Process in the Estimation of the Weight of the Landing Gear

#### 4.2.6. Payload Range Diagram

#### 4.3. Key Design Challenges

## 5. Airfoil and Wing Planform Selection

- The present eVTOL aircraft is taken to be flying at 2 km altitude in International Standard Atmosphere. The cruise speed is taken as 200 km/h.
- An aspect ratio of 8 was selected to reduce the span as much as possible. This will reduce the wing bending load due to placing the rotors on the wing. This will also enable the eVTOL to land on helipads.

#### 5.1. Airfoil Selection

_{max}at AOA 16 degrees.

#### 5.2. Estimation of Wing Parameters

#### 5.2.1. Angle of Incidence

#### 5.2.2. Twist Angle

#### 5.2.3. Three-Dimensional Lift Coefficient

_{tmax}is the wing sweep at the maximum thickness and a

_{0}is airfoil lift slope.

#### 5.2.4. Taper Ratio

## 6. Geometric Sizing

#### 6.1. Fuselage

- Seat width, 0.5 m;
- Aisle width, 0.5 m;
- Fuselage length, 6 m.

#### 6.2. Tail Sizing and Configuration Selection

_{HT}) is given by Equation (7).

_{HT}= √ {(2 V

_{HT}S

_{ref}c

_{ref})/(∏ (R

_{1}+ R

_{2}))}

_{HT}) was calculated using Equation (8),

_{HT}= (V

_{HT}S

_{ref}c

_{ref})/l

_{HT}

_{HT}) was calculated using Equation (9),

_{HT}= √ {AR

_{HT}S

_{HT}}

_{HT}is horizontal tail volume ratio, S

_{ref}is wing reference area, c

_{ref}is wing mean aerodynamic chord, AR

_{HT}is aspect ratio of horizontal tail, R

_{1}and R

_{2}are tail cone radius at the beginning and end, respectively. Since the tail is a V-tail, the values of S

_{HT}and S

_{VT}are chosen equal. The dihedral angle of the tail was selected as 45 degrees. The tail parameters are mentioned in Table 6.

#### 6.3. Powerplant Selection and Sizing

#### 6.4. Propeller Selection

- Number of blades in one engine = 3;
- Blade diameter = 1.50 m;
- Propeller collective pitch = 5°;
- Hub diameter = 0.40 m;
- Root twist = 32°; Airfoil used: NACA 64-935;
- Tip twist = 8°; Airfoil used: NACA 64-212.

#### 6.5. Landing Gear Sizing

- Ultimate landing load factor (N
_{L}) = limit load factor of gear (N_{gear}) × 1.5 = 20; - Extended length of main landing gear (L
_{m}) = 0.6 m; - Extended length of nose landing gear (L
_{n}) = 4 m;Main landing gear weight = 0.095 (N_{L}W)^{0.768}(L_{m}/12)^{0.409}Nose landing gear weight = 0.125 (20 × 1.5 × 1500)^{0.566}(L_{n}/12)^{0.845} - W
_{MLG}= 47.47 kg; - W
_{NLG}= 9.64 kg.

_{MLG}+ W

_{NLG}= 57.11 kg.

#### 6.6. Flight Controls/Avionics Sizing

_{Flight controls}= 0.053 L

^{1.536}B

_{w}

^{0.371}(N

_{z}W

_{dg}10

^{−4})

^{0.8}

- Fuselage structural length (L) = 8.23 m;
- Wing span (B
_{w}) = 10.97 m; - Weight of flight controls (W
_{flight controls}) = 26.23 kg.

#### 6.7. Cargo Mode

## 7. Design Parameters

#### 7.1. Constraint Analysis

^{2}and a power-to-weight ratio of 16 were selected. It can also be seen that the coefficient of lift required at the design point is 1.6 for a wing loading of 1000 N/m

^{2}.

#### 7.2. Sensitivity Analysis

- Maximum cruise speed, 73.352 m/s;
- Stall speed, 35 m/s;
- Maximum coefficient of lift, 1.6;
- Wing area, 17 m
^{2}.

## 8. Component Placement and Centre of Gravity Estimation

#### 8.1. Battery Location

#### 8.2. Electric Motor Placement

- To ensure that the crew and the passenger compartment are not within 5 degrees of the propeller disk in the event of a blade hurled towards the fuselage.
- The propeller must be as far as possible from the fuselage to avoid disturbance due to noise and aid in smooth flight.
- The span loading effect created will reduce the wing structural weight.

- The propeller hub and electric motor is able to rotate freely without any problems to facilitate various manoeuvres.
- The wing does not interfere with the thrust produced by the propeller during vertical flight.
- The fuselage drag is reduced by repositioning the fuselage away from the propeller’s wake region.

#### 8.3. Centre of Gravity (Initial Estimate)

_{n}, Y

_{n}and Z

_{n}for different components were obtained from OpenVSP. The already calculated weights W

_{n}, for components were used.

_{x}= (W

_{1}X

_{1}) + (W

_{2}X

_{2}) + (W

_{3}X

_{3}) + … + (W

_{n}X

_{n})

_{y}= (W

_{1}Y

_{1}) + (W

_{2}Y

_{2}) + (W

_{3}Y

_{3}) + … + (W

_{n}Y

_{n})

_{z}= (W

_{1}X

_{1}) + (W

_{2}X

_{2}) + (W

_{3}X

_{3}) + … + (W

_{n}Z

_{n})

_{cg}= M

_{x}/W

_{tot}

_{cg}= M

_{y}/W

_{tot}

_{cg}= M

_{z}/W

_{tot}

_{tot}= W

_{1}+ W

_{2}+ W

_{3}+ … + W

_{n}

_{cg}= 2.089 m, Y

_{cg}= 0 m and Z

_{cg}= 0.11 m. The CG shifts from X

_{cg}= 2.089 m at empty state to X

_{cg}= 2.21 m at maximum payload. Figure 18 shows the location of the CG in the OpenVSP view.

## 9. Aerodynamic Analysis

#### 9.1. Lift Curve

#### 9.2. Total Parasite Drag

_{total}= ∑ D

_{comp}+ D

_{L&p}+ D

_{misc}

_{comp}= D

_{SF}+ D

_{form}+ D

_{interference}

_{SF}= c

_{f}(S

_{wet/}S

_{ref})

_{form}= FF D

_{SF}

_{misc}is the drag due to miscellaneous items, D

_{L&p}is the drag due to leakages and protuberances, D

_{comp}is a component drag, D

_{SF}is skin friction drag, c

_{f}is coefficient of friction, S

_{wet}is wetted surface area and FF is the form factor. Drag contributions by several components of the eVTOL were tabulated in Table 11.

_{D}S

_{ref}= 3.83 u

^{2.5}A

_{max}

_{ref}is the wing reference area, q is dynamic pressure, u is upsweep angle and A

_{max}is maximum fuselage cross-sectional area [28].

#### 9.3. Calculation of Induced Drag

_{D}= C

_{D0}+ kC

_{L}

^{2}

_{D}is the coefficient of drag, C

_{D0}is the zero-lift drag coefficient, k is the drag factor due to lift and C

_{L}is the coefficient of lift [28].

#### 9.4. Lift Estimation

#### 9.5. Parasite Drag

#### 9.6. Total Drag

#### 9.7. Thrust Calculation

#### 9.8. Pressure Distributions

#### 9.9. Flow Interactions

^{2}/s at the nose of the fuselage and the minimum vorticity is −31.66 m

^{2}/s at the wing of the aircraft. The wingtip vortices get reduced by means of positioning the propeller by the tip of the wing.

## 10. Structural Analysis

#### 10.1. Structural Analysis of Fuselage and Wing Structural Members

#### 10.2. Material Selection

^{3}. The Ti 10-2-3 is a high-strength alloy and made up of titanium, aluminium, iron and vanadium. It has a tensile strength of 1170 MPa and a yield strength of 1105 MPa. Landing gear has to be developed in a certain way, where the forces will not surpass the maximum strength of 1105 MPa and it will not be damaged. Its melting point is 1649 °C and it has a high corrosion resistance [41].

## 11. Aircraft Performance

#### 11.1. Vertical Take-Off Performance

- Vertical take-off: The aircraft carries out a vertical take-off from 0 to 80 metres. It has no horizontal velocity. The vertical thrust is the outcome of the six engines of the aircraft pointed upwards. The thrust produced elevates the aircraft in the hover up phase. The vertical speed is 2 m/s.
- Climb: The eVTOL aircraft transits to a state where the wings also start to produce lift. This is conducted by controlling the RPM of the tail rotors to provide enough downforce of the tail to increase the angle of attack of the aircraft. From calculations, it was assumed that the wing will produce 75% of the lift required during the climb. The rest is produced by the propellers. It is also important to note that the propellers also overcome both the drag and weight during the climb. The climb phase takes place between the altitudes of 80 m to 2000 m. The rate of climb of the aircraft is 2 m/s.
- Cruise: The eVTOL aircraft completely transits to a state where all lift required is produced by only the wing. The propellers produce the thrust required for forward flight. The cruise phase takes place at an altitude of 2000 m for a range of 496.56 km. The mean horizontal speed was taken as the cruise speed, which is 69 m/s.
- Descent: The descent phase takes place from an altitude of 2000 m to 80 m at a descent rate of 2 m/s. The mean horizontal velocity is 21 m/s. This takes place for a horizontal distance of 1.72 km.
- Hover down: The hover down phase takes place from an altitude of 80 to 0 m. The hover down speed is 2 m/s.

#### 11.2. V-n Diagram

- Stall velocity, 35 m/s;
- Manoeuvring velocity, 68 m/s;
- Negative manoeuvring velocity, 55 m/s;
- Cruise velocity, 69 m/s;
- Maximum load factor n, +3.8;
- Maximum negative load factor n, −1.52.

## 12. Aircraft Stability and Control

#### 12.1. Longitudinal Static Stability

_{α}< 0). Various stability derivatives linked to lift and moments were estimated using the methods outlined in Raymer [28] and Anderson [44] and mentioned in Table 19. The moment coefficient at zero angle of attack is positive (Cm

_{0}> 0) and the trim angle of attack (α

_{e}) is well within the flight range, as shown in Figure 35. These two conditions indicate that the present eVTOL aircraft is both longitudinally stable and balanced.

_{Lα}is the lift curve slope of the wing. The negative value of Cm

_{α}, as shown in Table 19, shows that the present design has longitudinal static stability. An increase in the angle of attack resulted inactively in the creation of a restoring pitching moment. Moreover, the negative value of the derivative of the pitching moment coefficient with respect to the ruddervator angle (Cm

_{δe}), proposed that a positive ruddervator deflection further led to a counterbalancing pitch moment. The neutral point and static margin were found from the equations provided in the research of Raymer [28] and mentioned in Table 20. The lengths were measured from the nose of the aircraft, whereas the static margin was specified as a percentage of the mean aerodynamic chord. This positive static margin specified that the centre of gravity was way ahead of the neutral point, showing that this aircraft was longitudinal statically stable.

#### 12.2. Lateral-Directional Static Stability

_{β}shows that the aircraft certainly produced a restoring roll moment, when a change in the sideslip angle is given. The positive value of derivative Cn

_{β}specified that a counteracting yaw moment happened in the case of a sideslip change. Cl

_{r}and Cn

_{r}are derivative of rolling moment coefficient and derivative of yawing moment coefficient with respect to roll rate, respectively. These derivatives were calculated from the method outlined in Raymer [28] and were tabulated in Table 21, thus proving that the present design has lateral-directional static stability.

#### 12.3. Dynamic Stability

_{n}is natural frequency, ζ is damping ratio, T

_{P}is time period and T

_{1/2}is time to damp to half-amplitude.

#### 12.4. Dynamic Simulations

## 13. Cost Estimation and Technological Forecast

#### 13.1. Cost Estimation

#### 13.1.1. Cost Estimation Using Modified RAND DAPCA IV Model

_{E}) = 5.18 W

_{e}

^{0.777}V

^{0.894}Q

^{0.163}

_{T}) = 7.22 W

_{e}

^{0.777}V

^{0.696}Q

^{0.263}

_{M}) = 10.5 W

_{e}

^{0.777}V

^{0.484}Q

^{0.641}

_{Q}) = 0.133 (H

_{M})

_{D}) = 67.4 W

_{e}

^{0.63}V

^{1.3}

_{F}) = 1947 W

_{e}

^{0.325}V

^{0.822}FTA

^{1.21}

_{M}) = 31.2 W

_{e}

^{0.921}V

^{0.621}Q

^{0.799}

_{E}R

_{E}) + (H

_{T}R

_{T}) + (H

_{M}R

_{M}) + (H

_{Q}R

_{Q}) + C

_{D}+ C

_{F}+ C

_{M}+ Propulsion cost + Avionics cost

_{e}is empty weight, Q is lesser of production quantity or number to be produced in five years, V is maximum velocity, FTA is the number of flight-test aircraft, R

_{E}, R

_{T}, R

_{M}and R

_{Q}are average wrap rates of engineering, tooling, manufacturing and quality control. Thus, using the average costing method, the eVTOL aircraft for the prescribed design shall cost between the range of USD 14.83 M to USD 17.36 M.

#### 13.1.2. Cost Comparison with Existing eVTOL

#### 13.2. Technological Forecast in Battery and Cargo Systems

_{2}is emitted upon burning one ton of jet fuel and about 3.6 tons of CO

_{2}is emitted upon burning one ton of avgas. This clearly shows why using electric or at least hybrid power is very important. The usage of green propulsion is very important. Some of the specifications for an ideal battery include a low weight/volume ratio, high specific energy, high specific power, affordability, rechargeability, longevity, minimal toxicity and fast charging.

## 14. Relevance of Present Design and Analysis

## 15. Conclusions

## Author Contributions

## Funding

## Data Availability Statement

## Acknowledgments

## Conflicts of Interest

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Aircraft | Cruise Speed (km/h) | Maximum Speed (km/h) | Range (km) | Payload Weight (kg) | Power Plant | Crew | Propulsion |
---|---|---|---|---|---|---|---|

Vimana AAV | 244 | 280 | 900 | 400 | 200 KW Batteries | 4 | 8 propellers |

Vertical AerospaceVA-X4 | - | 241 | 161 | 450 | Batteries | 4 + 1 | 4 propellers for VTOL and all 8 propellers for cruise |

Rolls Royce | - | 402 | 804 | - | 500 KW Batteries | 5 | 6 propellers |

Overair Butterfly | 240 | 322 | 161 | - | Batteries | 4 + 1 | 4 propellers |

Hyundai S-A1 | 290 | - | 97 | - | 7 high density batteries | 4 + 1 | 4 propellers for VTOL and all 8 propellers for cruise |

Required | 200 | 250 | 500 | 500 | Green | 4 + 1 |

Mission Phase | Altitude | Horizontal Velocity | Vertical Velocity | Distance |
---|---|---|---|---|

Hover climb | mean sea level to 80 m | - | 2 m/s | - |

Transition and climb | 80 m to 400 m | 1.2 v_{stall} | 2 m/s to 1.2 v_{stall} | 0 to 120 m |

Accelerated climb (γ = 14°) | 400 m to 2000 m | 1.2 v_{stall} to 56 m/s | 8 m/s | 120 m to 1720 m |

Cruise | 2000 m | 56 to 69 m/s | - | 1720 m to 498.2 km |

Decelerated descent | 2000 m to 400 m | 56 m/s to 1.2 v_{stall} | 2 m/s | 498.2 km to 499.9 km |

Transition and descent | 400 m to 80 m | 1.2 v_{stall} | 2 m/s | 499.9 km to 500 km |

Hover descent | 80 m to mean sea level | - | 1 m/s | - |

Components | Weight (kg) |
---|---|

Fuselage | 240 |

Wing | 160 |

V-Tail | 26 |

Landing gear | 50 |

Battery | 513 |

Payload | 500 |

Motors | 120 |

Flight Controls | 26 |

Rotors | 20 |

Pilot/crew | 90 |

Other empty weight | 10 |

Total | 1755 |

Property | Dimensions |
---|---|

Span | 11 m |

Surface | 17 m^{2} |

Mean chord | 1.55 m |

Aspect ratio | 8 |

Taper | 0.35 |

Chord at root | 2.3 m |

Chord at tip | 0.8 m |

Twist | −2° |

Angle of incidence | 5.25° |

Sweep at quarter chord | 0 |

Mass | 200 kg |

Facing | Core | Adhesive | Finishing |
---|---|---|---|

Phenolic resin | Honeycomb-aramid fibre | Epoxy | Polyvinyl Chloride (PVC) |

Tail Parameters | Dimensions |
---|---|

Span | 4.439 m |

Taper ratio | 0.7 |

Aspect ratio | 3 |

Average chord | 1.48 m |

Dihedral angle | 45 degrees |

Velocity (v) m/s | Coefficient of Lift (C_{L}) | Coefficient of Drag (C_{D}) | Lift-to-Drag Ratio (L/D) | Thrust Required (T_{R}) N |
---|---|---|---|---|

70 | 0.634 | 0.047 | 13.6 | 1226.25 |

Type | Motor | Total Mass (kg) | Maximum Speed (RPM) | Torque (N-m) |
---|---|---|---|---|

BLDC | MP154120 | 20 | 9600 | 85.38 |

Tire Selection | Main Gear | Nose Gear |
---|---|---|

Outer diameter [m] | 0.34 | 0.33 |

Width [m] | 0.15 | 0.12 |

Rim diameter [m] | 0.101 | 0.101 |

Maximum load [kg] | 1600 | 800 |

Inflation pressure [kPa] | 930 | 792 |

Static loaded radius [m] | 0.13 | 0.13 |

Control Surfaces | Length (m) | Chord (m) |
---|---|---|

Ailerons | 2.1 | 0.2 |

Ruddervators | 1.99 | 0.2 |

Flaps | 1.5 | 0.345 |

Component | Coefficient of Parasitic Drag from Component Build-Up Method | OpenVSP Parasitic Drag Coefficient |
---|---|---|

Wing | 0.00965 | 0.0097 |

Fuselage | 0.007 | 0.006 |

V-Tail | 0.003 | 0.0038 |

Propeller Hubs | 0.0026 | 0.00264 |

Miscellaneous | 0.0137 | - |

Total | 0.0359 | 0.02214 |

Mode | Cruise mode |

Velocity | 70 m/s |

Pressure | 101,235 Pa |

Altitude | Sea level conditions |

Wing surface area | 16.8594 m^{2} |

Reynold’s number | 7.42 × 10^{6} |

Wing incidence angle | 5.25° |

Serial Number | Mode | Altitude (m) | Velocity (m/s) | Parasitic Drag (Cd_{0}) |
---|---|---|---|---|

1. | Cruise | Sea level | 70 | 0.03008 |

2. | Cruise | 2000 | 70 | 0.03085 |

Mode | Cruise mode |

Velocity | 70 m/s |

Projected area | 30.644 m^{2} |

Reynold’s number at sea level | 7.42 × 10^{6} |

Reynold’s number at 2000 m | 6.33 × 10^{6} |

Total Area | 4.4643 m^{2} |

Propeller Rpm | 6000 |

Air Velocity | 70 m/s |

Pressure | 101,235 Pa |

Density | 1.225 kg/m^{3} |

Altitude | Sea level conditions |

Mesh Elements | 121,135 |

Mesh Nodes | 22,016 |

Viscous Model | K-epsilon |

Serial Number | Load Type | Maximum Stress (GPa) | Maximum Deformation (m) |
---|---|---|---|

1. | Bending | 6.7 | 0.063 |

2. | Shear | 0.307 | 0.011 |

3. | Torsion | 6.587 | 0.227 |

Serial Number | Load Type | Maximum Stress (GPa) | Maximum Deformation (m) |
---|---|---|---|

1. | Bending | 32.6 | 0.092 |

2. | Shear | 0.198 | 0.001 |

3. | Torsion | 35.8 | 0.015 |

Mission Phase | Time (s) | Mean Horizontal Speed (m/s) | Mean Vertical Speed (m/s) | Energy (MJ) | L/D |
---|---|---|---|---|---|

Hover Up | 40 | 0 | 2 | 15.68 | - |

Climb | 80 | 21 | 2 | 70 | 12.3 |

Cruise | 7142 | 69 | 0 | 740 | 13.3 |

Descent | 80 | 21 | 2 | 70 | 12.3 |

Hover Down | 40 | 0 | 2 | 15.68 | - |

Derivatives (Unit—Per Radian) | C_{Lα} | Cm_{α} | Cm_{δe} | Cm_{0} |
---|---|---|---|---|

Values | 8.77 | −0.718 | −3.09 | 0.114 |

Parameter | Neutral Point | Static Margin | CG (Empty) | CG (Fully Loaded) |
---|---|---|---|---|

Values | 2.7 m | 29.4–41.2% | 2.0 m | 2.2 m |

Derivatives (Per Radian) | Cl_{β} | Cn_{β} | Cl_{r} | Cn_{r} |
---|---|---|---|---|

Values | −0.495 | 1.153 | 0.12 | −0.063 |

Dynamic Mode | Roots | ω_{n} | ζ | T_{P} | T_{1/2} |
---|---|---|---|---|---|

Phugoid | −0.015765 + 0.21582i | 0.217 Hz | 0.0728 | 29.1 s | 43.94 s |

Short period | −2.28323 + 3.6704i | 4.32 Hz | 0.528 | 1.71 s | 0.3 s |

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## Share and Cite

**MDPI and ACS Style**

Akash, A.; Raj, V.S.J.; Sushmitha, R.; Prateek, B.; Aditya, S.; Sreehari, V.M. Design and Analysis of VTOL Operated Intercity Electrical Vehicle for Urban Air Mobility. *Electronics* **2022**, *11*, 20.
https://doi.org/10.3390/electronics11010020

**AMA Style**

Akash A, Raj VSJ, Sushmitha R, Prateek B, Aditya S, Sreehari VM. Design and Analysis of VTOL Operated Intercity Electrical Vehicle for Urban Air Mobility. *Electronics*. 2022; 11(1):20.
https://doi.org/10.3390/electronics11010020

**Chicago/Turabian Style**

Akash, Arumugam, Vijayaraj Stephen Joseph Raj, Ramesh Sushmitha, Boga Prateek, Sankarasubramanian Aditya, and Veloorillom Madhavan Sreehari. 2022. "Design and Analysis of VTOL Operated Intercity Electrical Vehicle for Urban Air Mobility" *Electronics* 11, no. 1: 20.
https://doi.org/10.3390/electronics11010020