Research Article Volume 5 Issue 1
ELISA-Aerospace, France
Correspondence: Laurent Bovet, MINES ParisTech, Center for Materials Forming (CEMEF), UMR CNRS 7635, BP 207, 06904 Sophia-Antipolis, France
Received: November 25, 2020 | Published: January 28, 2021
Citation: Bovet L. An analytical computation of aircraft’s longitudinal trimmed conditions considering its fuel mass flow rate. Aeron Aero Open Access J. 2021;5(1):1-7. DOI: 10.15406/aaoaj.2021.05.00121
The aim of this article is to give simple expressions of aircraft’s longitudinal trimmed conditions taking into account the instantaneous fuel mass flow rate, i.e. the variation of total mass. The computation of aircraft trim points is not a new problem. Nevertheless, current analytical computations are classically performed with a constant total weight of the aircraft, hence assuming that the rate of decrease of the weight due to the fuel mass flow rate has insignificant effects on the results. Thus, the goal of this study has been to assess the effect of weight variation on aircraft trimmed condition and to compute correctly “extended trimmed conditions” defined as the equilibrated conditions in flight considering the weight decrease of the aircraft. It has been demonstrated that the weight variation must have an ad-hoc form to lead to extended trimmed conditions. Moreover, extended trimmed conditions do not correspond to a perfect level flight as it is the case at constant weight, but must present a slightly positive flight path angle leading to a regular increase in altitude. And the corrected longitudinal commands including throttle and elevator positions at a given airspeed and altitude have been computed for extended trimmed conditions and compared to the basic case at constant weight. Finally, all the analytical expressions given in this article have been verified through a numerical simulation performed in the case of a twin-engine aircraft representative of the Airbus wide-body family.
Keywords: longitudinal flight, trimmed conditions, fuel mass flow rate, flight dynamics
The computation of the aircraft steady-state or trimmed conditions is of importance in many flight dynamics problems. Indeed, trimmed conditions are the starting points to flight simulations1 and are used for further operations performed by pilots or engineers.2 Furthermore, linear models used for handling qualities analysis are derived from trim points,3 and performance analysis requires flight conditions to be as close as possible to the equilibrium.4 Finally, finding the general solution of trim states corresponding to given flight conditions can become non-trivial5 and lead to complex algorithms which aim at minimizing ad-hoc cost functions.6
Nevertheless, these analyses are always performed considering that the aircraft has a constant weight, assuming that the fuel mass flow rate–which is the airplane mass loss–has insignificant effects on trim. If this makes sense for short to medium term analyses, for long term, the weight decrease turns out in conflict with the initial trim assumptions. One of these concerns the well-known cruise at constant airspeed and angle of attack. In that case, the equilibrated states are computed assuming a constant altitude, but a simple analysis shows that the weight decrease automatically implies a gain of altitude and a residual flight path angle during the cruise.7
Indeed, a simple numerical computation of the longitudinal aircraft dynamics shows that the introduction of a weight decrease leads the aircraft to reach a slightly different equilibrated condition from the case computed at constant weight. Moreover, the well-known Bréguet equation predicting the range of an aircraft flying at constant airspeed and angle of attack has to be corrected by a small factor to take into account the mass variation in flight.8 Here this study makes a focus on the trimmed conditions taking into account the mass variation in flight and a simple way to compute these analytically by correcting the current flight dynamics expressions.
Condition on the state parameters
The equilibrium of a mechanical system is reached when the sum of the forces and moments are equal to zero. This leads to write that the acceleration of the body and the rate of change of angular momentum are null. Thus, considering a body with 6 Degree of Freedom (DoF) described by the 12 state parameters including:
The equilibrium is reached when the time derivative of each parameter is equal to zero, except for the 3 position coordinates for which the second time derivative is equal to zero. Thus, the 12 state parameters have been divided into two complementary vectors:
With this notation, the equilibrium is defined by:
Equilibrium⇒{˙→X=→0¨→Y=→0Equilibrium⇒⎧⎪⎨⎪⎩˙→X=→0¨→Y=→0 (1)
Case of a variable mass system
Let us assume that each coordinate xi of the vector →X→X is described by a first order differential equation gi function of the state variables and mass of the system. In addition, each position coordinate yi of the vector →Y→Y is described by a first order differential equation hi coming from the kinematics equations and function of the state variables only but not mass:
{˙xi=gi(x1,...,xn,y1,...,yk,m)˙yi=hi(x1,...,xn,y1,...,yk){˙xi=gi(x1,...,xn,y1,...,yk,m)˙yi=hi(x1,...,xn,y1,...,yk) (2)
The equations gi and hi are gathered in the vectors →G→G and →H→H resp., which leads to write:
{˙→X=→G˙→Y=→H⎧⎪⎨⎪⎩˙→X=→G˙→Y=→H (3)
Considering that a variation of mass affects the equations gi only, a variation dm during dt leads to a variation d˙xid˙xi and dxi given by:
d˙xi=∂gi∂x1dx1+...+∂gi∂xndxn+∂gi∂y1dy1+...+∂gi∂ykdyk+∂gi∂mdmd˙xi=∂gi∂x1dx1+...+∂gi∂xndxn+∂gi∂y1dy1+...+∂gi∂ykdyk+∂gi∂mdm
i.e.:
¨xi=∂gi∂x1˙x1+...+∂gi∂xn˙xn+∂gi∂y1˙y1+...+∂gi∂yk˙yk+∂gi∂m˙m¨xi=∂gi∂x1˙x1+...+∂gi∂xn˙xn+∂gi∂y1˙y1+...+∂gi∂yk˙yk+∂gi∂m˙m
which can also be written in the matrix form:
¨→X=Gx˙→X+Gy˙→Y+∂→G∂m˙m¨→X=Gx˙→X+Gy˙→Y+∂→G∂m˙m
with:
Gx=[∂→G∂x1;...∂→G∂xn]Gx=[∂→G∂x1;...∂→G∂xn]
and:
Gy=[∂→G∂y1;...∂→G∂yk]Gy=[∂→G∂y1;...∂→G∂yk]
The traditional equilibrium condition (EQ. (1)) gives in this case:
¨→X=Gy˙→Y+∂→G∂m˙m¨→X=Gy˙→Y+∂→G∂m˙m
This expression means that the equilibrated condition ˙→X=→0˙→X=→0 is generally not kept in time with a mass variation. Thus, another condition must be introduced to keep the equilibrium conditions (EQ. (1)) when the mass varies. This leads to the definition of an “extended trimmed conditions” given by:
Extended trimmed conditions⇒{˙→X=→0¨→Y=→0¨→X=→0Extended trimmed conditions⇒⎧⎪ ⎪ ⎪ ⎪ ⎪⎨⎪ ⎪ ⎪ ⎪ ⎪⎩˙→X=→0¨→Y=→0¨→X=→0 (4)
Considering that ¨→Y=→0¨→Y=→0 , the vectoris a constant noted →C→C . Then, (EQ. (4)) can also be written as:
Extended trimmed conditions⇒{˙→X=→G=→0¨→Y=→H=→CGy˙→Y+∂→G∂m˙m=→0Extended trimmed conditions⇒⎧⎪ ⎪ ⎪ ⎪ ⎪ ⎪⎨⎪ ⎪ ⎪ ⎪ ⎪ ⎪⎩˙→X=→G=→0¨→Y=→H=→CGy˙→Y+∂→G∂m˙m=→0 (5)
The theoretical framework of the definition of the extended equilibrium is now applied to the case of a pure longitudinal flight, i.e. a flight in the vertical plane only. In this case, the airplane has 3 DoF, i.e. 1 in rotation (pitch) and 2 in the vertical plane (horizontal and vertical movement), and 6 state variables describe the motion of the aircraft in this case.
Assumptions and models
In this application, the atmosphere state parameters are given by the International Standard Atmosphere (ISA) model.9 For analytical computations, an exponential model of air density is defined:10
ρ=ρrefeah(z−zref)ρ=ρrefeah(z−zref) (6)
where ρref and zref denote reference air density and altitude depending on the atmosphere layer.
In the troposphere, i.e. from sea level to 11km of altitude where the temperature decreases linearly, ah is approximated by a least-squares curve fit of the actual evolution of the air density vs. altitude, i.e.:
ah=−1/9042m−1ah=−1/9042m−1
In the lower stratosphere, i.e. from 11 km to 25 km of altitude where the temperature is constant, ah is directly
computed with the ideal gas law, i.e.:
ah=−1.577710−4m−1ah=−1.577710−4m−1
Drag and lift are the result of the projection of the aerodynamic force acting on the aircraft on the longitudinal
aerodynamic axes:11
{D=12ρSV2CDL=12ρSV2CL{D=12ρSV2CDL=12ρSV2CL (7)
In these expressions, V denotes the True Air Speed (TAS), S corresponds to the aircraft’s reference area and the coefficients CD and CL are the drag and lift coefficient resp. The drag coefficient CD is modeled as:15
CD=CDo+KC2LCD=CDo+KC2L (8)
where K corresponds the induced drag factor and CDo denotes the parasite drag coefficient.
Considering simple static models, the lift coefficient CL is a function of the angle of attack α only:
CL=CLα(α−αo)
Concerning the aerodynamic pitching moment M, we have:
M=12ρSV2cCm (9)
The pitching moment coefficient Cm is expressed in terms of angle of attack α, pitch rate q and elevator deflection δe:13
Cm=Cmo+Cmα(α−αo)+CmqqcV+Cmδeδe (10)
The thrust is assumed to be proportional to the air density and throttle lever δx:14
T=TSLρρSLδx (11)
In this expression, TSL denotes the total thrust of the aircraft at sea level.
Longitudinal equations of motion
As said above, it’s necessary to take into account 6 state variables to describe the motion of a pure longitudinal flight. The first 2 state variables concern the position of the aircraft in the vertical plane, i.e. the horizontal distance x and altitude z.
Each variable is given by a first order differential equation coming from the kinematics:
{˙x=Vcosγ˙z=Vsinγ (12)
where γ denotes the Flight Path Angle (FPA) of the aircraft.
With these equations, it is possible to define the vectors →Y and, i.e.:
→Y=(xz) and →H=(VcosγVsinγ)
Considering a symmetric aircraft with the thrust aligned with the fuselage reference line, the longitudinal equations of motion are the projections of the Newton’s second law on the aerodynamic axes:15
{˙V=1m(Tcosα−D−mgsinγ)˙γ=1mV(Tsinα+L−mgcosγ) (13)
Assuming that the thrust is located at the aircraft center of gravity, the pitch rate comes from the aerodynamic pitching moment only16 and we have:
˙q=MIYY (14)
This set is completed by the kinematic equations in pitch:
˙α=q−1mV(Tsinα+L−mgcosγ) (15)
With these equations, it’s possible to define the vectors →X and →G :
→X=(Vγαq)and →G=(1m(Fcosα−12ρSV2CD−mgsinγ)1mV(12ρSV2CL+Fsinα−mgcosγ)q−1mV(12ρSV2CL+Fsinα−mgcosγ)1IYY12ρSV2cCm)
Definition of the extended trimmed conditions
The first condition coming from the definition of the extended equilibrium (EQ. (5)), i.e. →G=→0 , leads to:
{Fcosα−12ρSV2CD=mgsinγ12ρSV2CL+Fsinα=mgcosγq=0Cm=0 (16)
The second condition from (EQ. (5)), i.e. ˙→Y=→C , gives:
{˙x=C1˙z=C2
where C1 and C2 denote the 2 coordinates of →C .
For the third condition from (EQ. (5)), it’s necessary to compute the matrix Gy:
Gy=[∂→G∂x;∂→G∂z]
i.e.:
Gy=[0ahm(Fcosα−12ρSV2CD)0ahmV(12ρSV2CL+Fsinα)0−ahmV(12ρSV2CL+Fsinα)0ahIYY12ρSV2cCm]
With (EQ. (16)), this leads to:
Gy=[0ahgsinγ0ahVgcosγ0−ahVgcosγ00]
We also have:
∂→G∂m=(−1m2(Fcosα−12ρSV2CD−mgsinγ)−gsinγm−1m2V(12ρSV2CL+Fsinα−mgcosγ)−gcosγmV1m2V(12ρSV2CL+Fsinα−mgcosγ)+gcosγmV0)
With (EQ. (16)), this leads to:
∂→G∂m=1mV(−gVsinγ−gcosγgcosγ0)
Finally, the third condition from (EQ. (5)) gives the following vector equation:
(ahgsinγahVgcosγ−ahVgcosγ0)˙z+1mV(−gVsinγ−gcosγgcosγ0)˙m=→0
This leads to a single condition on the evolution of the altitude in trimmed condition when the mass varies:
˙z=˙mah˙m (17)
Finally, the longitudinal trimmed condition of the aircraft taking into account the case ˙m≠0 is defined as follows:
Extended trimmed conditions⇒{˙V=0˙γ=0˙α=0˙q=0¨x=0˙z=˙mah˙m (18)
Discussion of the results
In extended trimmed conditions when the mass varies, the equations (EQ. (18)) show that the altitude must increase (˙z>0) when the mass decreases (˙m<0) to keep the equilibrium.
These equations also show that the condition on altitude ˙z=0 only apply to the case ˙m=0 , i.e. at constant mass. Moreover, as no equation in →G depends on the horizontal distance x, i.e. ∂→G∂x=→0 , it’s not possible to compute the constant C1 corresponding to ˙x , and the only condition in this case is ¨x=0 .
In addition, as we also have ˙z=C2 , this gives the following expression of ˙m :
˙m=C2ahm (19)
which means that the mass variationmust be proportional to the total mass m to get a trimmed condition in longitudinal flight. For an aircraft equipped with jet engines, this condition can be achieved through a very simple model with a fuel flow rate proportional to the thrust:
˙m=−CTT (20)
where cT denotes the thrust-specific fuel consumption (TSFC).
Indeed, considering the thrust given by the trimmed conditions, this leads to the following expression of the equilibrated thrust:
T=mg(sinγ+cosγCDCLcosα+sinαCDCL)
which gives:
˙m=−CTg(sinγ+cosγCDCLcosα+sinαCDCL)m
Assuming that the TSFC cT remains constant and as every parameter in this equation are constant in trimmed condition, including α and γ which are small, this leads to:
˙m=−CTgCDCLm (21)
Thus, this expression is equivalent to (EQ. (19)), which also gives the value of C2:
C2=˙z=−CTgahCDCL (22)
Analytical computation of the extended trimmed conditions
In this paragraph, the trimmed conditions in longitudinal flight at constant mass will be corrected to get the exact trimmed conditions when the mass varies.
To get this correction, the mass variation is assumed to be equivalent to the form given by (EQ. (19)), i.e.:
˙m=k˙mm (23)
with k˙m constant.
To clarify the notations, the subscript e denotes an equilibrated parameter computed with ˙m=0 , whereas the subscriptdenotes an equilibrated parameter computed with ˙m≠0 .
In addition, the comparison between both cases will be performed at a given weight for the same case of altitude and True Airspeed, i.e.:
{Ve=V˙mze=z˙m
As ˙z=˙mahm when the mass varies (EQ. (18)), by introducing both the kinematic equation in altitude (EQ. (12)) and the expression of m˙ (EQ. (23)), this leads to:
Veγ˙m=k˙mah
i.e.:
γ˙m=k˙mahVe (24)
Moreover, the pitch rate is equal to zero in both cases, i.e.:
qe=q˙m=0 (25)
The longitudinal trimmed equations are written at constant weight (lift and drag equations):
{Tecosαe−12ρeSV2eCDe=0Tesinαe+12ρeSV2eCLe=mg (26)
and with varying mass:
{T˙mcosα˙m−12ρ˙mSV2˙mCD˙m=mgsinγ˙mT˙msinα˙m+12ρ˙mSV2˙mCL˙m=mgcosγ˙m (27)
Assuming that the equilibrated FPA γ˙m is very small and considering that ρ˙m=ρe,˙V˙m=Ve and T˙m=Teδx˙mδxe , this
leads to:
{Teδx˙mδxecosα˙m−12ρ˙mSV2˙mCD˙m=mgγ˙mTeδx˙mδxesinα˙m+12ρ˙mSV2˙mCL˙m=mg
Assuming that αe and α˙m are small, the difference between each drag equation gives:
Te(δx˙mδxe−1)=mgγ˙m
which leads to:
εδx=CLeCDeγ˙m (28)
where ɛδx denotes the relative correction on the throttle lever δx to consider the mass variation, i.e.:
δx˙m=δxe(1+εδx)
Moreover, the difference between each lift equation gives:
Te(αe−δx˙mδxeα˙m)+12ρeSV2e(CLe−CL˙m)=0
which finally leads to:
εα=−εδx1+CLαCDe (29)
where ɛα is the relative correction on α to consider the mass variation, i.e.:
α˙m=αe(1+εα)
Considering that the longitudinal trimmed condition is also given by the condition Cm = 0 in both cases, we have:
{Cmo+Cmα(αe−αo)+Cmδeδee=0Cmo+Cmα(α˙m−αo)+Cmδeδe˙m=0
The difference between both equations gives:
Cmα(α˙m−αo)+Cmδe(δe˙m−δee)=0
which leads to:
εδe=−CmααeCmδeδeeεα (30)
where ɛδe denotes the relative correction on the elevator deflection δe to consider the mass variation, i.e.:
δe˙m=δee(1+εδe)
Numerical verification
These results have been verified through a numerical simulation performed in the case of a twin-engine aircraft representative of the Airbus wide-body family:17
The aircraft is assumed to fly in cruise at ze=30 000 ft and M=0.82, i.e. Ve=248.58 m/s, with=−10−5 kg/s which corresponds to a typical fuel consumption during the cruising phase. From initial conditions computed at ze and Ve, the mass variation ˙m=k˙mm has been applied abruptly, leading to the evolution of the longitudinal state vector computed with a numerical solver based on the Runge-Kutta algorithm.
Firstly, the computation is initialized with the trimmed condition with, i.e. γ0 = 0, α0 = αe, q0 = 0, δe = δee, δx = δxe. The figures (Figure 1) and (EQ. (2)) show that, after a transient including a classical “phugoïd mode” and a slow aperiodic mode, the TAS reaches a final state slightly different from the initial value. Moreover, the figure (Figure 2) shows that the FPA reaches a stabilized value γ˙m equal to its analytical expression, i.e..
γ˙m=k˙mahVe
The second step is to compute the evolution of the system with the simulation initiated with the corrected trimmed conditions with ˙m≠0 , i.e. γ0=k˙mahVe,α0=α˙m , q0=0, δe=δe˙m,δx=δx˙m . The figure (Figure 3) shows that the TAS (red curve) is well stabilized in time and almost equal to Ve in these conditions (relative difference equal to 4 10−6). Concerning the FPA, the figure (Figure 4) shows the transient in this case is very small and γ=γ˙m during the whole simulation.
This computation also shows that the correction ɛα and ɛδe can be considered as second order terms. Indeed, considering the simulation initiated with γ0=k˙mahVe , α0 = αe, q0 = 0, δe = δee, δx=δx˙m , i.e. without correcting the terms αe and δee, the evolution of the system (blue curve) is very close to the complete correction. In this case, the main difference lies in the transient which is slightly more pronounced.
Finally, the comparison between the trimmed conditions computed at a given altitude and airspeed considering a constant mass, i.e., vs. the trimmed conditions taking into account a mass variation given by ˙m=k˙mm , are gathered in the following tabular with the “first order” and “second order” corrections:
Correction |
First order |
Second order |
|
Flight Path Angle |
γe = 0 |
γ˙m=k˙mahVe | γ˙m=k˙mahVe |
AoA |
αe |
α˙m=αe | α˙m=αe(1+εα) |
Pitching rate |
qe = 0 |
q˙m=0 | q˙m=0 |
Elevator deflection |
δee |
δe˙m=δee | δe˙m=δee(1+εδe) |
Throttle position |
δxe |
δx˙m=(1+εδx)δxe | δx˙m=(1+εδx)δxe |
In this article, it has been demonstrated that the fuel flow rate of the aircraft affects the trimmed conditions. This has led to the definition of extended trimmed conditions, which are the equilibrated conditions in flight considering the weight decrease of the aircraft. When considering these extended trimmed conditions, correcting factors have to be applied on the control commands computed at constant weight to get the exact equilibrated flight taking into account the fuel mass flow rate of the aircraft. Moreover, it has been demonstrated that the mass derivative vs. time itself must be proportional to the total mass of the aircraft to get correct trimmed conditions. As a conclusion, the correcting factors given in this article give a sharper computation of trim points to be used as the starting points of numerical simulation and erase undue transients due to the sudden application of in-flight fuel burn. Moreover, these corrections should be taken into account to compute analytically the performance of the aircraft (endurance, climb, . . . ) as it has been already done for the range given by the Bréguet range formula.
None.
Author declares that there is no conflict of interest.
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