Vtol aircraft pdf

Create Alert Alert. Share This Paper. Background Citations. Methods Citations. Figures, Tables, and Topics from this paper. Citation Type. Has PDF. Publication Type. More Filters. Several concept vehicles have been developed, … Expand. The state of the art and operational scenarios for urban air mobility with unmanned aircraft.

The Aeronautical Journal. View 1 excerpt, cites background. Rotorcraft systems for urban air mobility: A reality check. Applied Sciences. Recent progress of electric systems has raised attention towards hybrid-electric and full-electric aircraft. This is probably not a good arrangement for most applications because the cockpit winds up in the rear for balance, which does not provide acceptable visibility for the pilot.

Also, in forward flight, the jet exhaust scrubs alongside the fuselage which causes thermal and acoustic problems. Tilt nacelles are heavy, but may be the best compromise for some applications. Grumman Aircraft Corporation has been pursuing a tilt-nacelle concept for Naval applications for a number of years. Some VTOL concepts provide a means of diverting the exhaust flow to provide vertical lift.

This is generally done by a retracting blocker device in the engine which shuts off the flow through the rearward-facing nozzle. The flow is then diverted forward through internal ducting. All of these VSTOL approaches, however, exact significant penalties in weight, cost, and completely when compared to a conventional jet aircraft design. In the area of propeller-powered aircraft, the tilt-rotor concept, as tested in the Bell XV, seems to offer the best compromise between helicopter-like vertical flight and efficient wing-borne cruise.

The tilt-rotor concept is the basis of the V Osprey. Subscribe Now! Sign In Sign Out. In contrast to using historical data, other studies have proposed to use momentum theory to calculate the power loading in designing tilt-wing and other transitional type of aircraft.

The first study 20 designed and tested a tilt-wing UAV, whereas the others to maintain flight 13 , 14 did the same process for a transitional UAV with two tilt-motors in the front, and a fixed coaxial ducted fan in the rear.

Unfortunately, the later UAV is also inherently unstable and when in operation, it heavily relies on the flight control system. Based on previous work, which has shown that it is possible to fly transitional tiltrotor aircraft in both modes VTOL and fixed-wing with only one aircraft configuration. Herein, we focus on the design of such aircraft with added lift surfaces, where most of the aircraft sub-systems are used in both flying modes.

Figure 1 Examples of tiltrotor, tilt-wing, and tail sitter UAVs. Figure 2 Simplistic designs of transitional quadcopters. Problem definition Obtaining stable and smooth transition in minimum time is one of the important requirements for TA, thus minimizing any operational losses that might be encountered in the transition phase.

Losing altitude during transition is one of the potential sources of instability that leads to a somewhat rough transition. In such case, the time needed to reach the wing-borne speed is increased. Under such conditions, the aircraft will lose altitude, stability and the transition time will increase resulting in performance deficiencies.

An alternative easy, but not optimal, solution is to use higher engine power. For these reasons, including the helicopter and transition flight modes in the initial sizing of TA is important to obtain the appropriate size of engine, wing and rotor. Due to the diverse complexities associated with diverse TA, this paper focuses on transitional UAVs that use the same propulsion system for both fixed-wing and helicopter flying modes, which is yet one of the most challenging configurations.

As most of the available UAVs use propeller-driven propulsion system, this paper will consider only tiltrotor UAVs with added lifting surfaces. Develop a preliminary design methodology for tiltrotor aircraft that would enable them to transition from VTOL to fixed-wing flight mode by reducing transition time, enhance stable transition and reduce dead weight for any specific flight mode while enabling close to optimal flight VTOL, Transition and fixed-wing which maximizes aircraft systems usage in all flying modes.

Due to the high complexities associated with the design of TA and the fact that the transition from VTOL to fixed-wing is of higher complexity compared to transition from fixed-wing to VTOL, the following assumption is considered: i it is assumed that the air density during transition does not change. That is, it is assumed that the altitude of the aircraft remains constant and that the range of speed is not large enough to have a change in air density surrounding the aircraft during the transition phase.

The reason for the transition from VTOL to fixed-wing being of higher complexity this is that the power required during transition from VTOL to fixed-wing is higher than the power required during transition from fixed-wing to VTOL.

Furthermore, due to potential large set of TA configurations and the scant available literature in the area and in order to further simplify the complexity of the problem at hand, some constraints are put in place when developing the proposed solution:. In order to address and provide a solution to the above-mentioned problem s under the assumptions and constraints described in Section 4, this paper proposes a preliminary design methodology for tiltrotor aircraft based on a three-step process.

This proposed three step process methodology comprise and takes into account the required three flying modes helicopter, transition and fixed-wing to generate a set of curves chart that will enable sizing the engine power loading , wing area wing loading , and the rotor disc area disc loading according to the required performance constraints during all flight modes. The suggested method is developed and targeted for aircraft that use the same engine for both fixed-wing and helicopter flying modes.

Due to the associated challenges, emphasis is placed on the transition from helicopter to fixed-wing whereas relaxing the operational flight conditions during the transition from fixed-wing to helicopter mode e. It is expected that with this proposed method, one can proceed to draw a set of curves that represent the required performance constraints in the three flight modes used in TA. As described in section 5, a mathematical approach is developed using variations of the typical mathematical formulations used in the independent design of fixed-wing aircraft and rotorcraft.

Such formulations are then combined to find a meeting point where both types of aircraft merge targeting tiltrotor TA. Sizing of the aircraft in helicopter flight mode is performed based on traditional rotorcraft performance equations 22 , 23 used within the aerospace industry. It is well known, however, that there are a large number of performance requirements that contribute to engine power and rotor size.

Example of such requirements include: the requirements for hovering at certain altitude, executing a vertical climb speed, hover at specific ceiling altitude, performing forward climb flight, performing some maneuvers, and descent flight. In order to simplify the process, the parameters that are expected to have the largest influence on the choice of the design point are first identified as suggested in, 22 , 23 and then used in the proposed procedure.

Thus far, the requirements for hovering flight, vertical climb flight, and ceiling altitude have been identified to have the higher power-to-weight ratio compared to forward climb flight and the descent flight. Additionally, such requirements are considered the most important in the choice of the corresponding rotorcraft disc loading.

So, the performance equations of these rotorcraft flight regimes are selected to be used in the sizing calculations. Generally, the total required power for rotorcraft to satisfy different performance requirements is given by the following equation: 22 , In the proposed analysis for tiltrotor UAVs, the tail rotor power and the transmission loss will be omitted e.

Additionally, depending on the flight mode, some of these power terms can be neglected if their values effects on the aircraft are very small compared to the others. During rotorcraft operation, a large number of transitional UAVs have the rotors over the wings that needed for fixed-wing operation e. This aspect complicates the design process, as extensive CFD has to be performed. However, if there is no wing area under the rotors i. From the one-dimensional axial momentum theory, 22 the induced velocity at the rotor in hover, v i , and the ideal induced power required to hover, P i , are given by Eq.

The obtained non-ideal induced power equation is then given by:. The rotor profile power P o , which is required to overcome the profile drag of the blades, is obtained by taking the drag of a blade element and integrating over the span of the blade Eq. In Eq. At this design stage, there are no precise data about a number of design parameters such as the number of rotors, rotor diameter, tip speed, airfoil type, solidity, drag, etc.

To account for the power needed to overcome these effects, the Figure of Merit FoM which is the ratio of ideal induced power for a rotor in hover obtained from momentum theory and the actual power consumed by the rotor Eq.

From Eq. The above equations are valid for configurations with no wing area under the rotor s. If this is not the case, the assumption that the rotor thrust equals the weight is no longer valid. This term varies from one aircraft configuration to another and depends on the equivalent drag area or drag coefficient per the wing to be used.

The second aspect in the proposed rotorcraft design aimed towards TA is to consider the required vertical climb flight. As there is no asymmetric flow in purely vertical climb on the rotor blades, tiltrotors act as a single main rotor helicopter. Thus, the basic requirements for vertical climb can be calculated based on the axial momentum theory.

Assuming there is no download force on the wing and neglecting the parasite drag of the aircraft, the climb and induced power required in vertical climb can be formulated as indicated in Eq.

Then, the total power needed can be formulated as Eq. Thus, reformulating Eq. In this approach, however, the service ceiling will be used in the performance equations instead of the absolute ceiling to allow hovering flight with a stability margin at the desired ceiling altitude. A typical value for the RoC at service ceiling for low subsonic aircraft is V y. Typical electric motors are designed to operate at sea-level ambient temperatures and below m.

Substituting Eq. Now, the identified parameters for the rotorcraft that will led to a TA have been obtained. The next phase of the proposed design approach that formulates the UAV preliminary sizing according to the transition phase requirements can be developed. These terms are selected because it is known that each of these power terms has a considerable value at moderate forward flight speeds obtained during transition.

These parameters are similar to the parameters used in the rotorcraft sizing equations presented in Section A. Therefore, it was concluded that it is possible to formulate the aircraft sizing for the transition mode base on modifying the sizing equations in rotorcraft mode and include the rotor tilt angle rather than modifying the sizing equations in fixed-wing and have them meet the rotorcraft sizing in the middle-transition phase as an alternative approach.

The reason for selecting this approach was due to the fact that it was identified to have more accurate calculation of the required power compared to omitting some power terms in fixed-wing approach e. Assuming it is required to size the aircraft to perform the transition at a fixed altitude, the power required to climb is zero. Assuming level transition flight flight path angle is zero , and for small angle of attack, the perpendicular component of forward flight velocity to the disc can be considered very small compared to the induced velocity at rotor disc.

So, the induced velocity can be approximated by momentum theory as given in Eq. According to the available rotorcraft forward flight performance equations, 22 , 23 the non-ideal induced power term during transition flight can be formulated as given by Eq.

Whereas the power required to overcome rotor profile drag is formulated as in Eq. Thus, per Eqs. Neglecting the small vertical component of the rotor profile drag, the force equilibrium in the vertical direction during level-flight transition can be expressed as a thrust force per Eq. Hence, reformulating Eq. This approach provides an effective approach that enables plotting the variation of the power loading with the corresponding disc loading at different values of wing loading.

Using the wing loading obtained from the fixed-wing performance formulations described in Section C , Eq. The last flight phase to be considered for a TA is its operation in fixed-wing mode. Similarly to the approach used when sizing the aircraft in rotorcraft mode, classical performance equations of fixed-wing aircraft 21 are herein used to allow obtaining a relationship function between wing loading and the corresponding power loading.

The four performance requirements selected in this paper for sizing the aircraft in fixed-wing mode are:. In addition, the requirement for the aircraft to perform both conventional and short runway Take-off must be considered and added to the approach as the aircraft might be required to perform such operation instead of VTOL to save power or be able to lift heavy loads, which might not be possible in VTOL.

Stall speed is one of the important fixed-wing aircraft performance requirements that limit the cruise speed to a minimum allowable value. It is important to consider such a limit in the design to prevent aircraft stall during its operation. The second considered performance requirement is the maximum forward speed. Both wing and power loadings contribute in achieving this parameter. As our concern is tiltrotor aircraft, the performance equations of propeller-driven aircraft are used to formulate a relationship between the power loading as a function of the wing loading and maximum speed as in Eq.

Another performance requirement is to climb in fixed-wing flight mode by a certain RoC. Similarly, the climb performance equations of propeller-driven aircraft are formulated to give the relationship between the power loading as a function of the wing loading and RoC as in Eq.

Ceiling in fixed-wing aircraft, the highest altitude that an aircraft can safely fly straight and level, is another requirement that affect wing and power loadings. In this formulation, we deal with the service ceiling instead of absolute ceiling to allow straight and level flight with a good stability margin at ceiling altitude.

The minimum take-off ground run distance is one of the important factors that affect the wing and power loadings. The take-off performance equations 21 of propeller-driven fixed-wing aircraft are modified to include thrust vectoring for Short Take-off and then used to derive the relationship between the power loading as a function of the wing loading and take-off distance as in Eq.

In Eqs.