Submitted byANSTINE MATHEW AUGUSTINE (32208101006) ARUN KRISHNAN. U (32208101009) MAHESH. J (32208101029) VETRI SELVAN. S (32208101057)
Military aircraft designed to attack ground and sea target by dropping bombs on them. Strategic bombers are designed for longrange bombing missions against strategic targets to damage enemy nations war effort .
Light bombers Medium bombers Dive bombers Fighters bomber Ground attack aircraft Multi role combat aircraft
Major type of aircraft designs Conceptual design Preliminary design Detailed design
It depends on what are the major factors for designing the aircraft. (a) Power plant Location: The Power plant must be located in the wings. (b) Selection of Engine: The engine should be selected according to the power required i.e., thrust required. (c) Wing selection: The selection of wing depends upon the selection of (1) Low wing (2) Mid wing (3) High wing - For a bomber the wing is mostly high wing configuration and anhedral. - Sweep may be required in order to reduce wave drag.
2. Preliminary design:
Preliminary is based upon number of factors like Loitering.
3. Detailed design:
In the detailed design considers each & every rivets, bolts, paints etc. In this design the connection & allocations are made.
To design a bomber aircraft Range of 20000 km & must carry 75000 kg+ of bombs & missiles. At supersonic & subsonic regimes To operate at regional bases with low cost of operation & maintenance The aircraft must also be capable of single pilot operation scenario. Due to long range pilot work load must be reduced The aircraft must be all weather , all terrain operation capable including the airbase. To take up a load factor +8g to +7.5g to -3.5g.
Collect data of existing aircraft of similar purpose i.e., bomber. The basic factors of aircrafts performance viz. Weight, Cruise velocity ,Range ,Wing area & Engine thrust. The performance data of various bomber aircraft with payload capacity between 5000 & 56600 kg was collected.
Mirage IIIE Mirage IVA F-111F F-111F swept Tu-22R Tu-85/1 YB-60 B-2A etc
From ComparisonParameters Max takeoff weight (kg) Thrust to weight ratio Aspect ratio Wing loading (N/sq.m) Span to height ratio Span to length ratio Combat radius (km) Pay load capacity (kmph) Max Speed (kmph) Service ceiling (m) Max Speed (m/s)
500000 0.28 8.4 7848 5 1.5 5000 75000 1000 15000 277.77
General rough estimate
Payload Fuel Structure Power plant Fixed equipments Total
Mass Fraction 0.15 0.45 0.32 0.07 0.01 1.00
Redefined Mass Estimation6 9000km 1000km
Mission profile for Strategic bombing
Analysis of mission profile TSFC values for BomberCruise 0.5 Loiter 0.4
Comparative data of Engines
Engine SelectionName of the Engine Manufacturer Type Length (m) Diameter (m) Wet weight (kg) Dry Weight (kg) Maximum Thrust (kN) Overall Pressure Ratio Thrust to Weight Ratio Fan Diameter (m) GP-7000 Engine Alliance Turbofan 2 Shaft 4.74 3.16 6800 6712 363 43.9 4.73 2.95
The above engine has been selected from a list of engines.
Redefined Thrust to weight ratio
content Airfoil nomenclature Lift coefficient Drag coefficient Types of airfoil Formula used Airfoil
AIRFOIL NOMENCLATUREThe cross-section shape obtained by the intersection of wing with the perpendicular plane is called airfoil. The major design feature of an airfoil is the mean chamber line ,which is the locus of points halfway between the upper and lower surface ,as measured perpendicular to mean chamber line itself . The most forward and rearward points of the mean chamber line are the leading and trailing edge respectively.
THE FORWARD AND REARWARD POINTS OF THE MEAN CAMBER LINE ARE THE LEADING AND TRAILING EDGES. CHORD LINE THE STRAIGHT LINE CONNECTING THE LEADING & TRAILING EDGES. MEAN CAMBER LINE THE LINE BETWEEN UPPER &LOWER SURFACES. CHAMBER MAXIMUM DISTANCE BETWEET THE MEAN CAMBER LINE & THE CHORD LINE.
LIFT COEFFICIENT The lift coefficient (CL or CZ) is a dimensionless coefficient that relates the lift generated by an aerodynamic body such as a wing or complete aircraft, the dynamic pressure of the fluid flow around the body, and a reference area associated with the body. It is also used to refer to the aerodynamic lift characteristics of a 2D airfoil section, whereby the reference "area" is taken as the airfoil chord. It may also be described as the ratio of lift pressure to dynamic
Drag Co-efficient:The drag coefficient (commonly denoted as Cd, Cx or Cw) is a dimensionless quantity that is used to quantify the drag or resistance of an object in a fluid environment such as air or water. It is used in the drag equation, where a lower drag coefficient indicates the object will have less aerodynamic or hydrodynamic drag. The drag coefficient is always associated with a particular surface area.
TYPES OF AIRFOIL CHAMBERED AIRFOIL SYMMETERICAL AIRFOIL
CHAMBERED AIRFOIL It is also called as unsymmetrical airfoil . Upper surface of the airfoil is not equal to lower surface. SYMMETRICAL AIRFOIL: Surface above the chord line and below the chord line are equal.
6. Airfoil selection and Wing Geometry estimates Main Parameter Selection: Wing Loading:
Thickness based Reynolds Number
Critical Mach number for the airfoil
LANDING GEAR TYRE SELECTION Load Distribution Typical load of aircraft while landing ;WL=WT -O.8WF While aborting mission ; WL=WT-O.1WF During static condition ; WL=WT
Ww=Ap x P Ap=2.3 dwww(dw/2-Rt) Rt=(dw/2-Ap/(2.3 dwww)) *RUN WAY LOADING Runway loading=load on each wheel/area of contact
DIMENSIONAL ESTIMATES Span to height ratio=b/ha 5 Span to length ratio=b/la 1.5 CONFIGARATION OF TAIL Horizontal stabilizer Horizontal stabilizer sizing 15% of wing area;sh/s=0.15 Vertical stabilizer geometry Vertical stabilizer sizing 9% of wing area ;sv/s=0.09
Configuration of tail
Airfoil NACA 0012
PREPARATION OF LAY OUT Wing location and C.G estimation Wfuselage X fuselage + W wing (X +X wing) = (Wfuselage+Wwing) (X+Xfinal) Where X is the location of wing root L.E from the nose fuselage and Xfinal is the reaction of cg from the L.E at root X final=0.35(Xcr - Xct)
Wing Detail for cg estimation
Three views of Aircraft
DRAG POLAR Drag equation for entire Aircraft:Cd=Cdwing+Cdothers+KCL^2 *wetted surface area Fuselage =Wfuselage*hfuselage Engine =4* /4d^2 Nose landing gear=dw*Ww*4 Main landing gear=dw*Ww*12 Main landing gear=dw*Ww*8 Flap=Lflap * Wflap
Take off performance =Cdpermanent+CdLG+Cdflap+Cdwing Landing performance=Cdpermanent+CdLG+Cdflap+Cdwing
Lift to Drag Ratio
Thrust required and Thrust available analysis: W1= 25% of Fuel and 100 % of Payload W1= 3185533.292 N W2= 50% of Fuel and 100 % of Payload W2= 3784962.23 N W3= 75% of Fuel and 100 % of Payload W3= 4384391.173 N
Thrust scenarios at Sea level for different weights
Thrust scenarios at 11 km altitude for different weights
Thrust scenarios at 25 km for different weights