Skip to main content

Full text of "THE DO 31 LANDING LOADS DURING VERTICAL LANDING AND THEIR CONSEQUENCES FOR FUTURE V/STOL DEVELOPMENTS"

See other formats


NASA TECHNICAL TRANSLATION NASA TT P-15,532 



THE DO 31 LANDING LOADS DURING VERTICAL LANDING 
AND THEIR CONSEQUENCES FOR FUTURE VSTOL DEVELOPMENTS 

W. Schoernack 



(KASfi-'XT-F- 15532) 'i?HE Do 31 LANDIHG H74-20568 

LOADS DUEING VESXICAL LAHDING AND THEIB 

CONSEQUEMCES FOE FUTURE V/STOL 

DEVELOPMEtSlS (Scientific Translation Duclas 

Service) 5$ P HC 45,75 CSCL OIC G3/02 36287 



Translation of "Die Landelasten der Do 31 bei 
Vertikallandungen und Folgerungen fiir zukiinftige 
VSTOL Entwlcklungen", Dornler-Werke G.m.b.H.^ 
Frledrlchshaften (W.Ger.), BMVg-FBWT-72-24; 

1972, 72 pp. 




NATIONAL AERONAUTICS AND SPACE ADMINISTRATION 
WASHINGTON, D. C. 205^6 APRIL 197^ 



NOTATION AND ABBREVIATIONS 



A 
A, B 



®x' \* 



f 
G 
g « 9,81 

s 

S/G 

t 

U 



Y 




iTiin 



I 



kp 1 Lift force 

Exponent (not defined more closely) 
Reaction force factors (e = 2P/G), referred 
to one undercarriage side 

Shock absorbing strut stroke 
Landing weight of the aircraft 

Acceleration of gravity 

Reaction forces which apply to one under- 
carriage unit 

Engine thrust In vertical direction 
Thrust/weight ratio 

s Time 

kn, it^/s Trajectory velocity 

rv/s Landlng-lslnklng velocity (at touchdown 

m/s Pictltjious landing sinking velocity (See p.l8j) 

Deg Trajectory angle relative to the horizontal 
Deg Longitudinal inclination angle of aircraft 
Deg Transverse inclination angle of aircraft 



m/s*i 
kp 



SUBSCRIPTS 



1 

2 

02 

left 

right 



First landing shock 
-Second landing shock 
Initial value at second touchdown 
Left main undercarriage unit 
Right main undercarriage unit 



ABBREVIATIONS 

CL 

SL 

VL 

CTOL 

STOL 

VTOL 



Conventional landing 

Short landing 

Vertical landing 

Conventional takeoff' and landing technique 

Short takeoff and landing 

Vertical takeoff and landing technique 



I^RECEDING PAGE BLANK NOT FILMED l|i 



TABLE OF CONTENTS 



Page 



I. LANDING LOADS OF THE DO 31 FOR VERTICAL LANDINGS 

AND CONSEQUENCES FOR FUTURE VSTOL DEVELOPMENTS 1 

1. Basic Aspects In the Design of Undercarriages 

for VSTOL Aircraft 1 

2. Design of the Do 31 and Test Results 4 

3. Undercarriage Design for Future VSTOL Aircraft 7 

II. EVALUATION AND RESULTS OP THE LANDING SHOCK MEASURE- 
MENTS 8 

1. Introduction 8 

2, Description of the Measured Variables 11 

2.1. Measurement of the landing shock forces 
and shock absorber stroke for the main 

main undercarriage 11 

2.2. Measurement of the sinking velocity and 
trajectory velocity of the aircraft 12 

2.3. Measurement of the acceleration 12 

2.4. Measurement installation 13 
3- Test Evaluation 13 

3.1. Evaluation of the landing shock forces 13 

3.2. Evaluation of the landing and sinking 

velocity 14 

4. Presentation of Results I7 

4.1. Tabular summary of the measurement 

results and evaluation results 17 

4.2. Graphical presentation of the measurement 
results and evaluation results 19 

5. Discussion and Evaluation of the Results 22 

6. Conclusions from the Results of the Landing 
Measurements 25 

TABLES 29 

FIGURES 32 

REFERENCES 5I 

ill 



THE DO 31 LANDING LOADS DURING VERTICAL LANDING 
AND THEIR CONSEQUENCES FOR FUTURE VSTOL DEVELOPMENTS 

Wolfgang Schoernack 

ABSTRACT: This report deals with the results of 83 ver- 
tical landings carried out during the Do 31 VSTOL Experi- 
mental Program. 

In 23 landings undercarriage reactions as well as sinking 
speeds were measured ^ of the remaining 60 landings only 
sinking speeds could be evaluated. Undercarriage reaction 
factors and sinking speeds are plotted as frequency distri- 
butions and are discussed. 

The result of the evaluation of the landing experiments can 
be summarized as follows: VTOL airplanes having a similar 
concept as the Do 31 and using manual control during the 
end of descent would experience considerably higher sinking 
speeds than conventional aircraft. It is remarkable that 
the frequency distribution of the sinking speeds is very 
severe, i.e., sinking speeds below 1 m/sec do not occur. 

Furthermore, a typical jumping of the airplane after 
touchdown and a following second impact prove unfavorable ^ 
this second impact resulting in higher undercarriage 
reactions than the first one. The horizontal loads 
occurring with vertical landings are smaller than expected. 

LANDING LOADS OF THE DO 31 FOR VERTICAL LANDINGS AND CONSE- 
QUENCES FOR FUTURE VSTOL DEVELOPMENTS 

1. Basic Aspects in the Design of Undercarriages for VSTOL 
Aircraft 



If one compares the landing techniques of conventional 
landing, short landing and vertical landing aircraft, we find 



/5J 



Numbers in the margin indicate pagination of original foreign 
text . 



that there are two characteristic parameters which characterize 
the state of the aircraft during landing, which are noticeably 
different for the three different aircraft categories: 

- The trajectory angle relative to the landing plane y 

- The trajectory velocity during the landing phase U 

Both state variables have a decisive Influence on the sinking 
velocity w perpendicular to the landing plane, which Is the most 
Important parameter for designing the undercarriage as far as 
shock absorption and strength are concerned. Therefore, they 
also Influence the design of the fuselage structure. 

The mutual dependence of the 3 parameters can be 
formally expressed as follows: 

s 

Y and U have the following opposing tendencies for the three /6 
aircraft categories : 

CTOL STOL VTOL 



Trajectory angle y Increasing 

Trajectory velocity U Decreasing 



Therefore, we cannot derive any tendency for the sinking 



velocity w from the above, 



Numerous civil and military specifications have evolved 
from experiences collected over many years. However, even the 
newest version only consider aircraft with conventional landing 
techniques and rotary wing aircraft as examples of VTOL aircraft 



The safe sinking velocities specified In these publications 
for structural design are as follows: 

- W » 2.15 (FAR 23) ... 3, OS (FAR 25) ... 4 m/s 0!II.-A-e862A) 
s 

for conventionally landing aircraft or, 

- Wjj " 2,0 (FAR 29) ... 2,5S m/s (PAR 27) 

for rotating wing aircraft.] 

Because there is no experience with STOL and VTOL aircraft, 
the prototype test facilities in the past had to specify landing 
parameters for experimental and prototype developments , which 
contained a certain safety margin. At the beginning of VSTOL 
development, a safe sinking velocity of w^ = 4.0 m/sec was 
assumed as a minimum, which was also used as a basis for the Do 31 
design. / 7 



In the meantime specification designs have been prepared 
in the United States and Great Britain over the last few years 
for STOL and VTOL aircraft, which already includes experience 
obtained with the first such aircraft. In the American design 
FARJ. XX, a safe sinking velocity of 2.6 m/sec is required for 
vertical landings (VL) and no numerical value is specified for 
short landings (SL). In the design of the British aircraft 
agency ARB, published with the title "Provisional Airworthiness 
Requirements for Civil Powered-Lift Aircraft", a minimum value 
of w = 2.15 m/sec Is required. Since there is no theoretically 
and experimentally based design value, a safe sinking velocity 
of w =4.6 m/sec is required, and no distinction is made between 
SL and VL. 

In other words there is no real agreement on these points. 
This is probably due to the fact that experience has been ob- 
tained in the United States with a prototype which is quite 



different from the VTOL aircraft used In Great Britain for this 
purpose . 

Another very Important aspect of the undercarriage design Is 
the question of the magnitude of the design side loads. For con- 
ventionally landing aircraft and most short landing alrcraf ts , 
the magnitude of the side load component during the landing 
shock depends primarily upon the sideslip angle and the oblique 
running properties of the tires. In the case of vertical landings, 
because of the slight side velocities, side loads will occur which 
are limited by the friction between the tires and the landing 
surface. In the most unfavorable case this means that the side 
load can have the magnitude of the vertical load. 

The construction specifications restrict the side loads of 
conventionally landing aircraft between 2^% (BCAR) and 40^ (FAR) 
of the maximum safe vertical loads, because of the limitations 
mentioned above. For rotary wing aircraft, FAR 27 and 29 
specify between 50 and 80S5 of the maximum reaction force as the 
design side force, depending on the loading case. The specifi- 
cation designs for VSTOL aircraft require 40^ of the maximum 
vertical loads for VL (FAR XX) or 50% for VL and 15% for SL /8^ 
(ARB suggestion). This means that there are different points of 
view here as well. 

2. Design of the Do 31 and Test Results 



The Do 31 was designed for short and vertical landings with 
a safe sinking velocity of 4 m/sec. At the time, this value was 
assumed to be required and also the NATO| specifications required 
this for a VSTOL fighter zone transport. 



Because there was no experience with VTOL aircraft, the 
American military specification for rotary wing aircraft MIL-S- 
8698 was used for the side force design at the request of the proto- 
type testing facility of the German air force (MBL) . However, the 
requirement for a 5055 side load in conjunction, with a vertical 
load corresponding to a sinking velocity of w^ = ^ m/sec would 
have led to weight Increases which could no longer be justified. 
Therefore, after agreement with the MBL had been reached, we 
defined a load case with w = 4 m/sec without a side load as well 
as a vertical landing case with w^ = 3 m/sec and a 50% side load. 

During the VTOL experimental program, 100 vertical landings 
were carried out with the Do 31-E3. 

The only accident which occurred involved the buckling of 
the main undercarriage legs and could be traced to a construction 
error of the locking mechanism. The permissible loading limits 
were not exceeded during this landing. 

Because of the harmless nature of this accident, we believe 
that we have a proof for the great safety of a VSTOL aircraft 
during the takeoff and landing phase. 

Only the last phase of the sinking flight is Important for 
the landing loads, that isjthe vertical descent from a height 
tietween four and five meters, as well as the touchdown process. 

The manual control of the touchdown process is made more 
difficult by the ground effects which occur (jet interference, 
recirculation), which leads to an increase in the sinking 
velocities. According to statements of the pilots, there was 
no way of Influencing the sinking velocity. 



/ 9 



The evaluation of the landing measurements shortly before 
contact with the ground therefore show a tendency for the sinking 
velocity to increase with decreasing height and a clear concen- 
tration at a value of 2 m/sec upon ground contact (Figure 17, l8) . 

The pilots were injstructed to turn off the main engines 
at the instant of ground contact in order to avoid a Jumping 
up of the aircraft. As experiments show (Figures 2, 3), it 
was never possible to avoid the rebound of the aircraft, probably 
because of the reaction times of the pilot and the engines. 
After the engine thrust had really dropped off, the second 
landing shock occurred with a strong loss in lift. The reaction 
forces were therefore larger during the second touchdown than 
for the first touchdown (Tables 2 and 3, as well as Figures H to 9) 
during most of the measured landings. 

Unfortunately, it was not possible to evaluate the sinking 
velocities after the first ground contact. Also, the rebound 
height as well as the thrust variation cannot be determined in 
practice. This means that only qualitative interpretations of 
the motion can be made during the second shock. / 10 

In the case of the vertical) landings of the Do 31, we always 
recorded side accelerations between 0.1 and 0.3 g and the 
corresponding side loads in the undercarriage. Therefore, we 
do not require a higher degree of safety than for conventional 
aircraft. 

The most important results of the Do 31 test can be 
summarized as follows, as far as the undercarriage is concerned: 



- No comparison with rotary wing aircraft can be made 

- Overall, the sinking velocities and the shock loads are 
extremely high compared with conventional landing 
techniques . 

- The safe sinking velocities must be assumed to be higher 
than for conventional aircraft. 

- The side loads which occur are not higher than for con- 
ventional aircraft. 

3. Undercarriage Design for Future VSTOL Aircraft 

The requirements on the undercarriage design of future 
VSTOL aircraft depend primarily on the sinking velocity changes 
caused by the ground effect and the possibilities of controlling 
them. 

In all configurations with a negative ground effect, the 
conditions will be similar to those of the Do 31, if the touch- 
down process is controlled manually. /ll 

Only a real VTOL aircraft can have a special design of the 
undercarriage shock absorbers, which would avoid or reduce the 
rebound of the aircraft and therefore avoid or reduce a second 
jump. In most cases it is likely that future VSTOL aircraft 
must be capable of short takeoffs and landings or conventional 
takeoff s and landings, especially if these are transport aircraft. 

Special touchdown automatic systems which would make it possi- 
ble to preselect the touchdown velocity could bring about basic 
Improvements in the magnitude of the safe sinking velocity, the 
shock. loads and influence the rebound upon ground contact. It 
seems that it will be necessary to develop this touchdown auto- 
matic system for civilian applications in particular. This is 
required not only because of the undercarriage design, but also 



It will result In simpler handling of the aircraft and higher 
passenger comfort. /12 

II. EVALUATION AMD RESULTS OF THE LANDING SHOCK MEASUREMENTS 

1. Introduction 

The design of the Do 31 provided an undercarriage arrangement 
with two main undercarriage units within the region of the two 
main engine gondolas at the wing, and a nose wheel unit. This 
arrangement led to relatively high main engine undercarriage 
strokes in conjunction with the high-wing configuration. 

Because of the special construction characteristics , 
limiting loads for the dimensioning of a large part of the wing 
structure for the load case resulted in "conventional horizontal! 
landing with wheel rotation shock" considering the dynamic 
amplification effects. This was based on the large torsion 
moments resulting from horizontal forces. In order to test 
the load assumptions and the structural calculations, especially 
for the dynamic loads and in order to provide experimental 
foundations for the calculation of such dynamic landing load 
cases, we installed measurement devices In the Do 31-El, the 
first experimental aircraft equipped for conventional flight 
testing. With this, the following quantities were recorded and 
measured as a function of time during l4 landings, in addition 
to the normal measurement program which includes physical data: 

- Forces in the main undercarriage in all coordinate 
directions . 

- Shock absorbing strut strokes. 



- Accelerations at the wing, fuselage, control surfaces 
and undercarriage. 

- Shear stresses at a- cross section of the inner wing. 

The qualitative evaluation of all conventional landings 
carried out led to the realisation that the load cases which 
had been assumed for the dimensioning could not be brought 
about during operation. The evaluation of two landings in [1] 
and [2] shows that the stresses caused by the undercarriage 
shock lie within the usual operational limits for conventional 
landings . 

During the VTOL test with a second experimental aircraft 
Do 31-E3i which was completely equipped with lifting engines, 
we found that considerably higher landing sinking velocities 
and therefore higher vertical undercarriage shock forces were 
achieved than during the conventional landings, in the case of 
vertical landings controlled manually by the pilot. As we will 
see, we particularly notice the large effect of the rebound of 
the aircraft which occurred regularly after first contact with 
the ground with a subsequent new touchdown. 

Since the pilot wants to reduce the rebound and wants to 
turn off the lifting engines as soon as possible after touch- 
down, in spite of the small sinking velocity, the shock force 
can be larger during the second touchdown than for the first 
touchdown. This is caused because the main thrust-weight ratio 
S/G is considerably smaller than one and in addition, the stroke 
has not yet been completely equalized during the second rebound. 

Since there were no measurement installations for determining 
the undercarriage shock forces in the Do 31-E3, we first only 
determined the landing sinking velocities from the radio height 



/13 



measurement In [i;]. This was then presented In the form of a 
frequency distribution. This evaluation already showed that the 
sinking velocity level and therefore the loads which occur during 
operation are considerably higher for vertical landings than for 
conventional landings. 

The measurement installation was expanded within the framework 
of collaboration with NASA on VTOL flight testing using the 
Do 3I-E3. It became possible to measure the loads in the main 
undercarriage during vertical landings. For this purpose, the A^ 
main undercarriage of the El was exchanged and built into the 
E3 and contained strain gauge bridges for determining the under- 
carriage loads. After adaptation of the measurement installation, 
we measured 23 vertical landings during the coujrse of the NASA 
test program (Figures 19 and 20). 

It is the purpose of the present report to obtain information 
on the undercarriage stresses of jet-supported VTOL aircraft 
derived from these measurements. 

In order to evaluate vertical shock forces during landing, 
we can state that the sinking velocity is the most important 
influencing parameter. The sinking velocity is difficult to 
measure and can only be measured inaccurately. Therefore, we 
use the force measurements to support and correct the measurement 
results. This was possible during the 23 landings for the reasons 
mentioned above, and for which we were able to obtain time 
variations of the reaction forces and side loads from the force 
measurements . 

It was possible to evaluate the sinking velocity measurements 
from sixty additional vertical landings. This means that the 
statistical data on the vertical reaction forces werej well founded, 

10 



so that they were included In the frequency distribution. 

In order to consider the number of load changes j which is 
important for fatigue calculations, the second landing shock 
had to be Included in the collection of sinking velocities by 
means of a flctit|ious sinking velocity. 

2. Description of the Me asured Variables .^ ^ 

L /15 

2_.1._ Measurement of_the_Land.ing_ShoGk_Porc^es_ and _ 

Shock Abs_orber_Stroke_ for the Maln_Unde_rc^arri_age_ 

The shock forces with components P^, P„ , P^ which apply at 
the main undercarriage units are referred to an aircraft-fixed 
axis system. )The positive directions of this system agree with 
the reference axes of the aircraft and are defined as follows 
(Figure 19) : 

x-direction: Positive backwards 
y-direction: Positive starboard 
z-dlrection: Positive upwards 

The shock forces apply at the wheel axis (P ^ P ) or at 
the tire contact surface (P ) and are measured using strain 
gauge bridges attached at suitable points of the undercarriage 
structure. They are calibrated. (Figure 20). The theoretical 
bases and the practical execution of the measurements (Skopinski 
method) are described in [1] and [5]- 

The shock absorber strokes are measured using potentiometers 
at the wheel control core. 



11 



2..2._ Measurement of_the_Slnklng_Veloclt;^ and TraJectory_ 
Veloc_lt_y_of the_ Alrcraft_ 

The sinking velocity is measured with an electrical variometer 
connected with the radio height measurement device. During the 
NASA test flights, a radio height measuring device with a 
measurement range between zero and 36O feet was installed for 
the lower height range. For the preceding test flights, a radio /I6 
height measuring device with a measurement range between zero and 
120 feet was installed. The measurement accuracy of the landing 
sinking velocity, which is influenced by ground roughness effects 
even for small horizontal velocities, is probably smaller during 
the NASA flights than during the preceding test flights. 

The horizontal velocity is measured using the Dornler- 
Pluglog. 

2_._3._ Measurement of_the_Ac_ce_leration _ 

The accelerations in the three axis directions are 
measured using the acceleration transducers in the vicinity of 
the center of gravity. These acceleration measurements could only 
be used to a limited extent for evaluating the shock process 
because the transducers did not respond to short time acceleration 
peaks, such as occur for the undercarriage shocks, at least in 
the z direction. Acceleration transducers installed in the shock 
absorbers for measuring the z-acceleratlons could only be 
applied infrequently within the framework of this evaluation. 
For the same reason, it was not possible to evaluate acceleration 
measurements in the outer gondolas. 



12 



2.4._ Mea.sureme_nt_ I^ns_tallat_ion_ 
The Do 31 measurement installation is described in [3]. /17 

The signals of the strain gauge bridges and acceleration 
transducers at the undercarriages or wing tips are recorded 
using the FM frequency multiplex system. The shock absorber 
stroke and the other flight measurement data of interest are 
measured using the time multiplex system. 



/ 18 



TEST EVALUATION 



3.1. Evaluation of the Landing Shock Forces 

The force components P^, P„ , P^ measured using strain gauge 

X y z 

bridges along the left and right main undercarriage sides are 
determined from analog recordings using the calibration coeffi- 
cients given in [5]. It was not possible to have a digital out- 
put of these measurement values because the digitalization pro- 
gram provided for this for frequency modulated measurement data 
could not be used. However, the evaluation of the analog record- 
ings was appropriate for the requirements, even though this 
required more time and even though the accuracy was lower. 

The components P (vertical force) and P (horizontal force) 
in the air craft /fixed coordinate system were determined from a 
calibrated strain gauge bridge, containing strain gauges in the 
the wheel axis. The P measurement bridge at the rear under- 

JPl. 

carriage dropped out after the first landing measurement. 

Since we did not obtain any useful results using the 
late measurement installation for measuring P^ and P^ 

a combination of three strain gauges, it was not possible to 

determine P at the : 

A. 

(Experiment No. 220) 



ultimate measurement installation for measuring P^ and P^ using 

possible 
determine P at the left undercarriage side, except for one landing 

A. 



13 



This latter statement is Inconsequential for the information 
obtained from these measurements, because the forces P^ are 
necessarily small for vertical landings with small forward 
velocities . 

The side force P was determined using the coefficients 
obtained from the calibration from the combination of two strain 
gauge bridges each. A comparison of the side forces with the /19 
associated side accelerations showed that apparently the sign 
of the y force of the left undercarriage was wrong because of 
an error in the installation of the measurement installation. 
This error was corrected when the results were evaluated and 
processed. 

Prom the analog recordings of P it is obvious that the 

aircraft and the main undercarriage rebounds after the first 

touchdown from the ground in all of the 23 measured landings. 

The shock forces P during the second touchdown are greater 

z 

than for the first touchdown for all of these landings. For this 
reason, the undercarriage shock forces were determined separately 
for the first and second landing shock. 

3-2._ Evaluation_of the Landing and Slnklng_ Velo£lty_ /20 

Prom the measurement data of the radio height measuring device] 
or the connected electrical variometer, we find sinking velocity 
values which fluctuate more or less and sometimes increase in 
the vicinity of the ground. The following influences could be 
responsible for these phenomena: 

- The signals of the radio height measuring device designed 
for the 360 foot measurement range are disturbed by 
ground roughness effects during the vertical landings 
carried out with relatively small forward velocities. 

14 



This can result In a considerable scatter of the signals 
towards the ground because of the large measurement range 
of the radio height measuring device. 

- The engine thrusts did not remain constant during the 
descent. Because of hot gas circulation and Jet inter- 
ference, they can decrease in an irregular fashion in the 
ground effect zone and cannot be controlled by the pilot 
in order to obtain a uniform, unaccelerated final descent. 
This means that In this region, the main thrust-weight 
ratio S/G is already smaller than 1 and the sinking 
velocity Increases. 

This means that the determination of the landing sinking 

velocity w from this radio height measurement contains some 
s 

degree of uncertainty. In contrast to this, the measurement of 

the landing shock forces P„ , P^ and their determination from analog 

y ^ 
recordings can be considered to be relatively accurate. In order 

to support the experimental determination of the landing sinking 
velocity w , we carried out motion calculations of the landing 
process on the computer in parallel with the test evaluation. 
We used the known design data of the undercarriage system and 
determined the shock forces P as a function of the sinking 
velocity w , and we assumed a thrust-weight ratio of S/G=l, as 
well as the average values of the true landing weights and 
landing inclination angles. Figure 1 shows P^ plotted as a /21 
function of w for G = I85OO kp and 6 = 3° as obtained from 
this calculation. The calculation applies for a symmetric land- 
ing, i.e. for equal shock loads on the left and right sides of 
the undercarriage. 

Since the shock forces P^ are not the same on the left and 
right sides for most of the measured landings, the true resistance 

15 



parallel to the ground cannot be taken Into account in the calcu- 
lation and because S/G Is already smaller than one for the first 
touchdown, of course the determination of the landing sinking 
velocity of the aircraft from the measured shock forces P^ will 
be inaccurate using the Impact calculations mentioned above. 

The approximate determination of the sinking velocity w^ 
for the first touchdown from measured impact forces P^ accord- 
ing to the method given above can be looked upon as adequate 
for a statistical evaluation of the measurement results in the 
form of frequency distributions. 

As already mentioned, the analog traces of P^, show that 
the aircraft rebounds after touchdown and lifts away from the 
ground during all 23 measured landings. 

For the subsequent second touchdown, the analog recordings 
have a much flatter increase of P^ which decreases over the shock 
time, compared with those of the first touchdown. Since the 
shock forces P are greater compared to the measured sinking 
velocities than for the first touchdown, from this we may con- 
clude that the lifting thrust-weight ratio S/G is much smaller 
for the second landing shock than for the first one. Prom this 
it follows that the lifting engines are turned off more or less 
rapidly by the pilot after the first ground contact. Since /22 
there are no sufficiently differentiated lifting thrust data 
available for the landing phase, it is not possible to determine 
the most important parameters S/G besides the sinking velocity 
w during the second landing shock within the framework of this 
paper, which could then be used for a statistical frequency 
evaluation. During the second landing shock we find that the 
shocks during the second landing shock have not been completely 
extended by the springs when the second touchdown occurs, accord- 
ing to measurements of the shock absorbing struts. This is 

16 



expressed by an Increase in the shock load. Pour exceptions to 
this were measured. 

Even though the sinking velocity w ^i during the second land- 

S ii| 

Ing shock alone does not give any information on the shock forces 
if S/G is also not known, we nevertheless determined the values of 
w„T from the sinking velocity measurement using the electrical 
variometer connected to the radio height measurement device. 
This was done in order to have a means of comparison. However ^ 
it Is not possible to check or correct the measured values of 
w 5 as before using the measured shock forces when w -, is deter- 
mined. This is because the magnitude and variation of S/G are 
not available for the second landing shock. 

In order to obtain a characteristic frequency representation 
by means of the sinking velocity alone during the second landing 
shock especially for the Do 31-E3, flctit[ious computered landing 
sinking velocities w ^ were determined uslng|the shock calculation. 
This calculation satisfies the requirement that at S/G = 1 and 
for initially completely extended shock absorbing strokes, the 
measured shock forces P ^ during the second landing shock are 
approximately obtained, and are therefore comparable to w -, for 

S J- 

the first landing shock. 

4. FRESEMTATIOW OF RESULTS /23 

4.1. Tabular Summary of the Measurement Results and 
Evaluation Results 

Table 1 shows the following landing state variables for 
evaluating the shock forces: 



17 



- Landing weight G 

- Trajectory velocity U 

- Longitudinal inclination angle 6 

- Transverse Inclination angle ^ 

- Sink velocity for the first touchdown w ^ 

- sink velocity for the second touchdown w ^ S-'^^ ^s?' 
respectively. 

The values of G are taken from the available weight summaries. 
The values of U,9 and (f> are taken at the time of the firstj 
touchdown from the "quick look" data. The sinking velocity values 
w , J w „ and w „ are determined using the approximate method 
described before in Section 3.2. 

Table 2 shows the undercarriage shock forces P^-, , P -, , P -, andl 

xl y 1 ' zl [ ' 

the shock force factors e , , e , , e , determined from them 

xl ' yl zl 

(referred to G/2) as well as the shock absorbing strut strokes, 
f-j at the left and right main undercarriage side during the first 
landing shock. 

In the same way. Table 3 shows the shock forces P^n^ ^vP' ^7?* 

which result during the second landing shock, the shock factors 

e ^, e ^, e „ as well as the shock absorbing strut stroke f ^ at 
x2y2z2 o2 

the time of the second touchdown and the maximum value fp during 
the second landing shock. In experiment No. 23^, we did not ob- 
tain any measured values for determining the shock absorbing strut 
strokes. The values of f-^p in Table 3 show that for the landings 
considered, except for experiment No. 240, and on one side each 
for experiments Wo. 24l and 244, the shock absorbing strut was 
not completely extended at the time of the second touchdown. 



18 



The values of P , P and P assumed In Tables 2 and 3 are ^^ii 
always maximum values, which do not always exactly coincide in 
time. 

4.2. Graphical Presentation of the Measurement Res ults and 
Evaluation Results 

Figures 2 and 3 show the time variation of the shock forces 
P and P of two typical landings. Figure 2 shows experiment 
No. 228 with the greatest P^ force during the first landing 
shock. Figure 3 shows experiment Mo. 243 with the largest P^ 
force during the second landing shock. There is a typical P^ 
variation for all 23 measured landings. Clearly we can see the 
lift off of the aircraft from the ground after the first landing 
shock and the subsequent second developed landing shock. 

In order to compare the stresses of the Do 31 undercarriage 

according to P and P during the first and second landing shock 

encountered during the 23 measured vertical landings. Figures 

4 to 9 show frequency distributions of the shock force factors 

e and e ^ with a classification interval of Ae =0.3- Figures 

al z2 
10 and 11 show frequency distributions of the shock force factors 

e ^ and e „ with a classification interval of Ae = 0.05- The 

y 1 y 2 <> 

representations show that the operational stresses as measured 

by P and P are harder for the second landing shock than for 

the first landing shock. The maximum values of P^ and P^ do not 

always occur at the same time. L /^^ 

Since the shock force P^ in an undercarriage depends 
primarily on the landing sinking velocity w^ for a constant S/G, 
the operational load values of P^ in general will be determined 
through the frequency distribution of w^ . This is why we 
determined comparable frequency distributions of w^ using the 



19 



values of P determined from experimental measurements-: 

In Figure 12 we show the frequency distribution of the 

landing sinking velocity w^, for the first landing shock 

obtained from 23 vertical landings with a classification 

interval of Aw =0.5 m/sec. In the same way, we show the fre- 

s _ , 

quency of the fictitious values w for the second landing shock | 

determined far S/G=l from a numerical comparison.) 

Figure l4 gives In the usual manner the frequency of exceed- 
ing w (during the first landing shock) per flight, as determined 

S -L 

from 23 landings. Considering the fact that during these landings, 
two well developed landing shocks occurred (as though two landings 
were flown one after another for each flight), we find the exceed- 
ing frequency per flight given in Figure 15 for a flctitjious landing 

sinking velocity w , I assuming a thrust-weight ratio of S/G = 1 

s 

and the frequency distribution of w^^ according to Figure 12 and the 
frequency distribution of w , according to Figure 13. 

Since the statistical Information content of a frequency 
evaluation of the landing sinking velocity is relatively small 
as obtained from 23 flights, we included the frequency distribu- 
tion of the landing sinking velocity w^-, determined earlier in 
[4] from 60 vertical landings. In order to have a more compre- 
hensive evaluation. These 60 vertical landings were carried out 
during the Do 31-E3-VISTOL basic testing programs. However, 
this extended evaluation was restricted to the landing sinking 
velocity w -, (first touchdown), because for these landings, no 
undercarriage shock forces were measured. Consequently we do 
not have any experimental data on the magnitude or existence of 
a second well-developed touchdown shock during these landings. 



20 



Figure 16| shows the frequency distribution of the landing 
sinking velocity w •, was determined [4] from 60 vertical landings! / 26 

S -L 

during the Do 31-E3-VSTOL basic testing program. These values 
were determined graphically from the radio height measurement 
(measurement range from - 120 feet). 

From the combination of the frequency distributions of 
w -| according to Figure 12 (from 23 landings during the NASA 
test program) and Figure l6 (from 60 landings of the basic 
testing program) we find the frequency distribution of w , from 
a total of 83 vertical landings of the Do 31-E3 as shown in 
Figure 17. The frequency distributions w -, in Figures 12 and 16 
show that the landing sinking velocity level is higher on the 
average during the 23 landings carried out during the NASA test 
program than for the 60 landings carried out during the basic 
testing program. It should be noted that during the 60 landings 
performed during the basic testing program, there was no check 
of the sinking velocity values determined from the radio height 
measurement by means of shock force measurements. 

Figure I8 shows the exceeding frequency w , per|fllght 

S J- 

which results from the frequency distribution w -, according to 
Figure 17 and from 83 vertical landings with the Do 3I-E3. 
As a comparison we also show the exceeding frequencies according 
to the American military specifications MIL-A-8866 for conventional 
aircraft landing on airports. The dashed lines refer to training 
aircraft and the dash and dot lines correspond to normal military 
aircraft of other types. ThereforeJPigure I8 shows that the 
sinking velocity level during the vertical landings of the 
Do 31-E3 is ] considerably higher than for conventional landings. 
This Is shown even more clearly in Figure 15, because for example 
during the 23 vertical landings measured during the NASA test 
program^ there were two well developed landing shocks which occurred 
per flight because of the rebound after the first touchdown shock. - 

21 



In most of the 23 landings which could be evaluated' in this regard, 
the shock forces P ^ and therefore the comparable computed sinking , 
velocity w ^ were larger than for the first landing shock. 

S i— 



5. DISCUSSION AND EVALUATION OF THE RESULTS 

One important result is the fact that for all 23 measured 

vertical landings, because the aircraft rebounded after the first 

touchdown, there was a second well developed landing shock and the 

shock forces P are greater than for the first landing shock for 
z 

most of these landings. (Tables 2 and 3, Figures 4 to 9). The 
number of asymmetric landing shocks , for which P^ are not the same 
on the left and the right undercarriage sides, is greater for the 
second touchdown than for the first touchdown (Tables 2 and 3). 
The true sinking velocity w ^ during the second touchdown is 
usually considerably smaller than for the first touchdown com- 
pared to the shock force P^ (Tables 1, 2 and 3). The large P^ 
forces during the second landing shock are conditioned by the fact 
that the power turns off the lifting engines as soon as possible 
after the first touchdown in order to avoid the rebound. This 
means that S/G during the second landing shock is substantially 
smaller than one. In addition, the shock absorbing struts are 
usually not completely extended when the second touchdown occurs, 
(fo2 in Table 3). 

The fictitjious sinking velocity values w^j ^°^ '^^^ second 

touchdown which are comparable with the sinking velocity values 

w , which occur during the first touchdown, which were determined 

si 
approximately from the shock forces P^ using shock calculations 

for S/a = 1 and initially completely deployed shock absorbing 

struts, are larger during the corresponding landings than the 

sinking velocity w -, during the first touchdown (Table 1), Just 

like the shock forces P . The rebound of the aircraft after the 

first touchdown shock and for the same undercarriage essentially 



22 



depends on the landing sinking velocity w . For the same sinking 
velocity, rebound Is Just as possible during conventional landings 
with A/G = 1 as during vertical landings with S/G = 1. 

The reasons forjthe large rebound of the aircraft after /28 
touchdown are therefore the high landing sinking velocities 
w -, which In our case are considerably higher than during con- 
ventional landings (Figures l4 and l8) . 

The operational stresses on the aircraft caused by the shock 
forces P are| represented by the frequency distributions of the 
landing sinking velocities (Figures 12, 13 and 1?) and the exceeding 
frequencies per flight (Figures 1^1, 15 and l8). These stresses are 
much greater for vertical landings of the Do 31-E3 than for the 
conventional landings. This is easily seen by the increased| 
sinking velocity level for the first touchdown shock (Figures 
l4 and l8) and is amplified by the second touchdown shock (Figure 
15). 

The largest landing sinking velocity determined during a\ 

total of 83 landings is w = 3-4 m/sec and is therefore 15^ 

s 

smaller than the largest design sinking velocity of w = 4 m/sec, 
which was substituted for the Do 31 for the design landing weight 
of G = 21800 Kp. 

The largest shock force P on one main undercarriage side 

determined from measurements of 23 vertical landings is the same 

for the first and second landing shocks andlamounts to P , = P ^ = 

' zl z2 

18000 Kp (Tables 2 and 3). This means that it is about l8^ 

smaller than the shock force P^ - 22000 Kp assumed for the 

design case "Two-point horizontal landing with 25^ drag" 

assumed to occur at w =4 m/sec and G = 21800 Kp . 

s 



23 



The largest shock absorhlng strut stroke determined from 22 
vertical landings and which of course occurs during the second 
landing shock is fj = 337 mm (Table 3) and Is therefore 19^ 
smaller than the maximum possible stroke of HlS mm. 

The shock forces P are greater for most of the 23 measured 
vertical landings than for the first landing shock (Tables 2 and 
3), I jjust like the forces P^. /29. 

However, their relative magnitudes nor their time of ) 
occurrence are usually not related to those of the P^ forces. 

The curves showing the frequency distributions of the maximum 
absolute values of the shock factors e^^ during the first landing 
shock in Figure 10 and those of 62 ^ov the second landing shock 
in Figure 11 clarify the harder frequency distribution of P^| 
during the second landing shock. 

The largest shock force P on one undercarriage side is 

P = 4000 kp for the first landing shock (Table 2) and P p = 
yi 
3800 kp for the second landing shock (Table 3). The largest p^ 

forces do not coincide with the largest P^ forces (which can be 

explained by the fact that the magnitude of P^ depends primarily 

on the side motion of the aircraft (sideslip landing). 

In the design assumptions for the DO 31, we assumed according 
to the helicopter specifications MIL-S-8698 and for a vertical 
landing with a side wind of w^ = 3 m/sec and S/G - 2/3, a side 
force of P = 0.5-P . With a design landing weight of G = 21 
800kp, we have P^ - 19 ^00 kp and therefore P^ or 6^=2 'P^/G =0.9, 
respectively, per main undercarriage side. 

Since the magnitude of the P forces probably depends 
greatly on the operational and landing characteristics (for 

2k 



example,! side wind and airport roughnesses), the results obtained 
from the 23 vertical landings of the NASA testing program are not 
sufficient for generally valid statements on permissible side 
loads. This is especially true if we consider the especially 
favorable environmental conditions (good weather conditions, 
selected landing sites). 

The P forces could only be determined on the right under- 
carriage side during the 23 measured vertical landings, with 
the exception of one landing (experiment No. 220), because the 
corresponding strain gauge brldge]failed. As expected, they 
are smaller because of the small horizontal landing velocity, 
compared with the P„ forces which are decisive for the design. 
These forces occur for the load cases (braking) , "vertical 
landing spin up" and "two-point landing fuselage down spring- 
back" for conventional landings. - /3Q 

This means that the P forces determined from these measure- 
ments are very restricted and do not suffice for a general 
statement regarding the expected P forces during vertical land- 
ings. 

6. CONCLUSIONS FROM THE RESULTS OF THE LANDING MEASUREMENTS 

Considering certain restrictions such as the relatively 
low number of measured landings and the fact that the testing 
was performed by experienced test pilots with very extensive /31 
safety measures, we can nevertheless draw some conclusions 
for the design of VTOL aircraft. We can derive some inter- 
actions to be used in the conception of new landing techniques 
and methods. The configuration and design of the engine installa- 
tion of the Do 31 represents another factor which limits the 
generality of our conclusions. Nevertheless, we believe that the 
basic conclusions drawn should be valid for other engine 

25 



configurations . 

The vertical landings performed by the Do 31 were manually 
controlled according to the pilot's vision during the last phase 
of descent (from a height of 5-10 m) . The test results show 
an accumulation of the measured sinking velocities at a value of 
2 ra/sec with a maximum value up to 3.4 m/sec. Sinking velocities 
below 1 m/sec do not occur at all. Compared with conventional 
landings, vertical landings are in general much harder according 
to the sinking velocities (Figures l4 and l8). 

Another result of the test is that there is a rebound of 
the aircraft after the first touchdown which results in a second 
shock with considerably higher shock loads. It is included in the 
collection of sinking velocities by introducing a fictitjious 
sinking velocity. This then leads to an even harder collection of 
sinking velocities (Figure 15). Considering the design sinking 
velocity of 4 m/sec for the Do 31, this means that there is a 
smaller safety margin compared with conventional applications. 
This is especially true considering the much more unfavorable 
fatigue loads caused by the landing shock forces. 

Based on the present results we may conclude the following: /32 
If the minimum requirements for landing sinking velocity are to 
be applied for future developments as specified in the specifi- 
cations "FAR. XX" of the PAA or the "Provisional Airworthiness 
Requirements for Civil PL-Aircraft" of the ARB, this can only 
be done under the assumption that another landing technique is 
used. For example, a sinking velocity control could be used 
which would assure that the sinking velocities specified in the 
design documents would not be exceeded during operation. Such 
a landing technique would be especially advantageous for 
civilian passenger traffic, especially from the point of view 
of passenger comfort. These requirements should be easy to 

26 



comply with because a sinking velocity control system Is 
necessary any way In order to control the vertical landing. 

In addition, a positive Influence on the rebound after 
touchdown can be obtained by optimlElng the undercarriage shock 
absorber system in the direction of vertical landings. Also 
the shock loads could be reduced in this way. This would also be 
advantageous for the design of a control system. 

We cannot give any statistical information on the measured 
side loads P because of the small number of measured landings. 
Compared with the vertical loads, the measured side loads lie 
within the design limits. In this connection it is Interesting 
to note that according to Information obtained from the pilots , 
the side motions of the aircraft are easy to control during 
hovering flight. Therefore, we do not expect any special 
difficulties, at least for flight over flat land. 



Civilian Regulations 



/3^ 



USA: 



Federal Aviation Regulations (PAR) 



PART 23 
PART 25 
PART 27 
PART 29 
PART XX 



Valid for small aircraft 

Valid for commercial aircraft 

Valid for small helicopters 

Valid for small commercial rotary aircraft 

Design of VSTOL aircraft 



27 



Great Britain: 

British Civil Airworthiness Requirements (BCAR) 
Section D: Airplanes: valid for aircraft In general 

Provisional Airworthiness Requirements for Civil Powered-Llft 
Aircraft: design for VSTOL aircraft 

Military •Specifications: 



USA; 

Military Specification MIL-A-0088f2A 

"Airplane Strenght and Rigidity, Landing and Ground 

Handling Loads" 

Military Specification MIL-A-008866A 
"Airplane Strength and Rigidity RellabilitV 
Requirements, Repeated Loads and Fatigue" 

Military Specification MIL-S-8698 

'Structural Design Requirements, Helicopters" 



28 



TABLE I. LANDING STATE VARIABLES 



/35 



'Exp. 
No. 


Q 


U 

[icu] 


e 

(deg) 


* 

. (deg) 






|m/t«c) 


220 


t424S 


G 


7 


0,5 


2,3 


U 


1.9 


222 


18 325 


20 


4 





3.0 


1,4 


2.8 


223 


17 805 


20 


7 


2 


2,7 


1,1 


2,7 


225 


17 827 


2 


8 


3 


1,9 


U 


2,2 


326 


18 230 


22 


6 


-1.5 


2,4 


1.4 


2.6 


227 


18 270 


9 


9 


1 


1.9 


0.8 


1,9 . 


223 


17 605 


Ifi 


8 


0,5 


3,4 


1,2 


2,5 


229 


18 320 


16 , 


5 


-1,6 


2,6 


1,5 


3.0 


231 


17515 


» 


6 


-0,5 


3,1 


1,4 


3.2 


232 


18 403 


25 


6 


t 


3,9 


U 


2,5 


233 


18 275 


19 


3 


1 


2,6 


t,6 


2.9 


234 


18 703 


13 


8 


2 


2,7 


1.2 


2,0 


235 


18»3S 


10 


13 


3,6 


1.9 


1,4 


2.6 


236 


18 700 


S 


7 


0,5 


2.2 


1.6 


2.B 


237 


18137 


4 


6 


-I 


2,1 


1.8 


2.3 


238 b 


18 330 


3 


8 


-2,6 


t,6 


1.4 


2,3 


239 


18400 


20 


a 


1.6 


2;> 


1.2 


2.8 


240 


18£<>8 


31 


8 


-4 


1,9 


3.1 


3,3 


241 


18132 


26 


8 


-1,6 


1.9 


1,4 


2,6 


243 


18645 


»> 





2 


2.4 


2.3 


3,4 


244 


18 700 


12 


? 


1 


1.7 


1,6 


*•« 


247 


tasse 


M 


S 


2.5 


2.3 


1.6 


3,1 


24a 


18S33 


\2 


s 


2.B 


2.2 


U 


2.4 



29 



TABLE II. UNDERCARRIAGE SHOCK FORCES, SHOCK 

FORCE FACTORS AND SHOCK ABSORBER STRUT 
STROKES FOR FIRST LANDING SHOCK 



/36 



i 


Exp A 
No./ 


Left 

Right 




Left 

Right 


P.. (kp( 

Left 
Right 


Left 

Right 


Left' 
Right 




Left 
Right 




f, (mm) 

Left 
Right 




, 


!M 


- bW 

- 460 


740 
1 390 


9bO0 . 
11 600 


-0,031 

-0,05 


o.cai 

0.152 


1.04 
1,26 


241 
269 




122 


370 


- 700 
7M 


14 GOO 
14 000 


0,04 


- 0.07$ 
0.079 


1,E3 
1.53 


305 
304 


■ 


223 


1 

- E60 


520 
1070 


11600 
13500 


-0.OG3 


0,059 
0.12 


1.29 
1.52 


260 
305 




225 


-T 110 


-1 to 

3030 


6 200 

10 400 


-0,125 


-0,115 
0.345 


0,69 
1,15 


139 
217 




226 
227 


-2 WO 


-2 070 

-2 0GO 


100C0 
11 500 


- 0.224 


- 0.227 
-0.223 


1,1 

1.2S 


223 
271 




-1300 


- 650 
6P0 


9500 
7EiM 


-0,142 


-0.071 
0,074 


1,04 
0,C3 


233 

173- 






22C 


-2 590 


2010 
2 240 


1800Q 

15CO0 


-0,234 


0,228 
0,255 


2,04 
1.71 


319 
313 - 






729 


-1G70 


1 100 ■ 
1 &40 


12 COO 
12CO0 


- 0,182 


0.12 
0.179 


1,31- 
1,31 


2G0 
250 






231 


- 1 850 


2D50 
2360 


Hfioa 


-0.211 


-0,233 
0,27 


1,69 
1,69 


313 ■ 

3ca 




232 


-2S30 


2000 
1690 


14 000 
13 600 


-0,28 


0,216 
0.203 


1,51 
1,47 


301 

2ca 




233 


-4 480 


1 730 
1 9S0 


11 400 . 
11600 


- 0.162 


0,19 
0,216 


1,25 
1,27 


270 
273 




234 


-2040 


650 
1650 


12 600 

13 000 


-0.218 


0,091 
0.177 


1,35 
1.39 


282 
287 




23S 


-2400 


370 
1940 


7 800 
9 20O 


-0,254 


0.04 
0,205 


0,82 
0.97 


197 
226 




236 


-1 480 


-2900 
-1 880 


9000 
•T500 


-0,158 


-0,31 
-0,2 


0.96 
1.12 


-** 




237 


- 930 


- 670 
-1 930 


oSOO 

9 200 


-0,102 


-0,053 
-0,213 


1,03 
1.02 


248 
241 




738b 


740 


-1370 
-2 350 


8000 
5000 


- 0,078 


-0,145 
- 0,238 


0.B5 
0.53 


193 
114 




239 


2220 


600 
960 


9B00 
11 200 


0,24 


o.oss 

0.105 


1,07 
1.2 


235 
262 




2tO 


- 930 


-3C80 
300 


9500 
6000 


-0,1 


-0,332 
0,032 


1,02 
0.65 


218 
161 




241 


-2410 


1270 
300 


7200 
9200 


-0,256 


0,14 

0,033 


0.79 
1,02 


165 
227 




243 


- 830 


2 320 
3290 


lOBfO 
11600 


-0,1 


0.25 
0.3 


1.17 

1,25 


266 
261 




344 


-1 110 


920 
620 


8000 
6800 


-oTl2' 


0,099 
0.066, 


. 0.G6 
0,62 


219 
1S5 




247 


1200 


2040 
4000 


10 000 
11000 


0,13 


0,222 
0,43« 


1,09 
1,2 


223 
240 




24« 


- 370 


1210 
1420 


9 800 
10E00 


,t0.m 


0.129 
0.151 


1,04 


249 

-.Z» 



30 



TABLE III. UNDERCARRIAGE SHOCK FORCES, SHOCK FORCE 

FACTORS AND SHOCK ABSORBING STRUT STROKES 
FOR SECOND LANDING SHOCK 



/37 



'Exp. 

No. j 


P., [kpl 

Left 
Right 


P,, Ikpl 

Left 1 
Right 


Left 
Right 


'Left 1 
Right 


Left 
Right 


Left 
Right 


Left 
Right 


\ |mmi 

'Left\ 
Right^ 


220 


13S0 

1390 


-940 
-2 300 


6000 
9 600 


C,152 
0,152 


-0.103 
0,252 


c,r,6 

1,05 


lib 

155 


ite 

254 


222 


- 370 


-2&G0 
-2 (WO 


11000 
13000 


- 0,039 


-0,28 
-0.31 


1,2 
1,43 


120 
200 


277 
297 


223 


1950 


3 050 

3 800 


1340O 
12000 


0,219 


0.343 

0,427 


1,51 
1.35 


117 
134 


295 
249 


225 


1 200 


3 800 
2 870 


11 000 
8400 


0,134 


0.424 

0.32 


1.23 

0,04 


102 
107 


265 
2-16 


226 


650 


-1 720 

-2 60O 


11 400 

12400 


0,071 


-0.1D9 
- 0.285 


'.25 
1,36 


62 
132 


277 

290 


227 


370 


680 
- 910 


9000 
8 000 


0,039 


0.QO& 
-0,099 


0.9it 
0,83 


100 
107 


227 
230 


228 


740 


-2 860 
-2?60 


10000 
12200 


0,OS4 


- 0,325 

- 0,336 


1,14 
1,39 


152 
187 


255 
301 


229 


- "370 


980 
- 840 


14 000 
15000 


-0.04 


0,107 
-0,092 


1,53 
1.64 


12S 
120 


304 
3?0 


231 


- 740 


-3 180 
-2 350 


12 200 

16 600 


- 0.085 


- 0,363 
-0,268 


1,39 
1,92 


152 
172 


235 
337 


232 


370 


-1 B50 

- 1,990 


11 000 
12600 


0,01 


-0,2 

-0,215 


1,19 

1,35 


179 
157 


277 

?90 


233 


- 740 


BOO 
- 900 


13 000 

14 000 


-0,081 


0,088 
-0,099 


1,42 
1.G3 


135 
124 


297 
312 


234 


660 


- I 000 
1 640 


9 200 
9000 


0,06 


- 0,107 
0.175 


0.93 
0,96 


86 
63 


248 
2-13 


235 


-2 600 


2 960 
1 110 


13 000 

10 200 


- 0.275 


0,3 
0,117 


1.33 
1.03 


95 
94 


295 
255 


236 


- 190 


2 060 
1600 


14 600 
9 COO 


-0,02 


0,22 
0.17 


1,56 
1,03 


^ 





237 


- 190 


-2 070 

-2 530 


9 000 
11 200 


-0,02 


- 0,228 

- 0.285 


0,99 
1,24 


80 
93 


240 
272 


238 b 


- 650 


-3 690 
- 1690 


7 200' 
12 800 


-0,07 


-0,39 
-0.18 


0,76 

1,35 


84 
57 


223 
242 


239 


190 


- 820 
740 


uooo 

13 200 


0,02 


-0,09 
0.08 


1,41 
1,44 


103 
139 


295 

301 


240 


1200 


i690 
-2 850 


.16 000 
17000 


0,13 


0,29 
-0,31 


1,73 
1.B3 






304 
3J9 


241 


- SCO 


-2 050 

2 930 


12 400 

11300 


-0,062 


- 0,226 
0,323 


1.37 
1.24 



82 


230 
263 


243 


1 200 


1 370 
1030 


16 000 
10000 


0,13 


0,148 
0,111 


1,73 
1,94 


87 
78 


319 
322 


244 


1 110 


-2 470 
-2 300 


13 500 
1VOO0 


0.12 


-0,264 
- 0,246 


1,44 


72 



238 
, 281 


,247 


- 460 


1090 
700 


13 800 
16000 


-0,05 


0,119 
0,076 


1.5 
1,74 


68 
70 


306 
321 


34B 


- 740 


440 
960 


9 800 
12000 


- 0.079 


0.047 
0,102 


1,04 
1.28 


104 
120 


263 
385 



31 




Figure 1. Shock force P^ as a function of sinking velocity w^ 
(from motion calculations of the landing shock process) for 
Do 31-E3I 



32 



-'Main undercarriage leftinj 



Main undercarriage rightfti. 

vi.i:. I lit; 




Figure 2. Measured time variation of the landing shock forces 



P and P from Experiment No. 22 8 



33 




/40 



Measured time variation of the landing shock forces 



Figure 3' 
P and P from Experiment No. 213 
z y 



34 



/4l 



u 
P4 



50 . 


■ 


















40 . 


■ 


















30 . 


■ . 












• 






20 . 
























V 


10 
















1 — j^ 


1 — ^ — i 
















' 






1 — j__j 


' 





0,4 0,7 1 1,3 1,6 1,9 

Shock force factor e , left! 

zl 



2,2 



Figure 4. Frequency distribution of the shock force factors 



e ^ left from 23 vertical landings 



35 



/42 



o 


cr 

u 

fLt 



40 . 


. 










' 








m- 








30 . 


■ 










fl 




. 












20 . 


. 




10 . 


• 








» 






0|4 


0,7 


1 


1,3 


1#6 


1,9 



Shock force factor e - right 



Figure 5, Frequency distribution of the shock force factor e^-^ 
rlKht from 23 vertical landings 



36 



/ns 



1>^ 
o 

c 

01 



50.. 



40,, 



30.. 



20.. 



10.. 



■■ ■!■ - 

0,4 



0,7 



1,3 1,€ 1,9 



Shock force factor e . left 

zl 



Figure 6. Frequency distribution of the shock force factor e 
left from 23 vertical landings 



zl 



37 



/44 



50 .. 



40 .. 



o 

c 

QJ 

1-1 

P4 



30 .. 



20.. 



10.. 



0,4 0,7 I 1,3 1,6 1,9 

Shock force factor e ^ right \ 



Figure 7. Frequency distribution of the shock force factor e^-j^ 
right from 23 vertical landings 



38 



/45 



50 .. 



40 .. 






30 .. 



20 .. 



10 .. 



■4- 



+ 



H- 



■+■ 



0,7 1 1*3 1'^ I'S 

Shock force factor e , 

zl 



2,2 



Figure 8. Frequency distribution of the largest shock force 
factors e -, left or right from 23 vertical landings 



39] 



50 .. 



40 .. 



o 

t-l 
P4 



30 .. 



20.. 



10.. 



/46 



0,7 



1,3 1,6 1,9 2/2 



Shock force 
factor e 



-z2 



Figure 9- Frequency distribution of the largest shock force 
factors e^2 ^^^^ o^ right from 23 vertical landings 



40| 



30 .. 



/M7 



o 

a 

0) 

3 

01 
U 



20 .. 



10 



■+■ 



+ 



■4- 



0,05 Orl 0,15 0,2 0,2S 0,3 

Shock force factor e 



0,35 0,4 0,45 



yi 



Figure 10. Frequency distrllDutlon of the largest shock force 
factors e -, left or right from 2 3 vertical landings 



41 



/il8 



30 .. 



a) 

cr 
0) 



20 .. 



10.. 



-t — f. 



H *- 



0,05 0,1 0,15 0,2 0,2S 0,3 0,35 0,4 0,45 

Shock force factor e ^ 



Figure 11. 
factors 



'y2 



Frequency distribution of the largest shock force 
leftlor right from 23 vertical landings 



42' 



/49 



40 .. 



30 ., 



o 
w 



20 .. 



10 .. 



1 1,5 2 2,5 3 3,5 

■Landing sinking velocity w ,1 



m 

sec 



Figure 12. Frequency distribution of the landing sinking velocity 
w ^ from 23 vertical landings during the NASA test program 



41 



/50 



60 .. 



5Q .. 



40 






3 J .. 



20 ,. 



10 



1,5 2 2,5 3 3,5 

' - - - Landing sinking velocity] — 

I "s2 

>n_ j 
. sec ' 

Figure 13- Frequency distribution of the fictitfious landing 
sinking velocity w from 23 vertical landings during the NASA 
test program 



44 




Figure 14. Exceeding frequency of the landing sinking velocity w -, per 
flight from 23 vertical landings during the NASA test program 



vn 













Figure 15. Exceeding frequency of the fictitious landing sinking velocity w^ 
per flight including the second landing shock from 23 vertical landings 
during the NASA test program 






/53 



40 .. 



30 .. 



20 .. 



10.. 



-f- 



-♦- 



+ 



■+■ 



-+- 



0,1 1 1/5 2 2,5 3 

Landing sinking velocity w 



3.5 



si 



m 



sec 



Figure 16. Frequency distribution of the landing sinking 
velocity w , according to [4] from 60 vertical landings for 

basic test program 



47| 



40. 



/54 



30.. 






2C.. 



iO.. 



0,5 



H — 
1,5 2 2,5 3 3,5 

Landing sinking velocity w 

m 
sec 



Figure 1?. Frequency distribution of the landing sinking ■ 
velocity w ^ from 83 vertical landings (23 landings during NASA 

test program and 60 landings for basic training program) 



48 




Figure 18. Exceeding frequency of the landing sinking velocity (8) per flight (*) 



MD * Translator's note: Illegible in foreign text. 



U1 



T6e<t> 




Figure 19 



Ground line of the compressed" tl re^|. 
Arrangement and design of main undercarriage 



/56 



50 



131 



Measurement points on the main undercarriage o£ the Do 31 El| 




BOtO($l)Z 



Plffure 20. Strain gauges BD on the main undercarriage 



51f 



REFERENCES 



1. Calculation of Dynamic Load Cases. Dornier Report, ZTL Task, 

No. 53-822-K-ii+l. 

2. Further Development of Modern Structure Calculation Methods. 

Dornier Report, ZTL Task, No. TO135/9^100/924l2 . 

3. Development of the Transport Aircraft Do 31 and Production 

and Testing of Two Experimental Aircraft. Dornier Report, 
GE 51-481/69 

Summary Result Report on the Development Period from 
1962 to October 31, 1969 . 

4. Evaluation of the Landing Sinking Velocity from Flight 

Measurements for Vertical Landings of the Do 31-E3. 
Dornier Report, 1503I-O2. IO56 

5. Do 3I-EI Measurement of Dynamic Loads on the Nose and 

Main Undercarriage using Strain Gauges. Dornier Report, 
EV31-292. 



Translated for National Aeronautics and Space Administration 
under contract No. NASw 2483, by SCITRAN, P. 0. Box 5456, Santa 
Barbara, California, 93108. 



52'' 



STANDARD TITLE PAGE 



1. Report Ha. 

NASA TT F- 15.532 



2. Government Aeeenion No. 



4. Title end Subtitle 

THE DO 31 LANDING LOADS DURING VERTICAL LANDING 
AND THEIR CONSEQUENCES FOR FUTURE VSTOL DEVELOP- 
MENT S_____^ . '. ■-—- 



7. Aulhor(i) 

W. Schoemack 



3. Recipient'* Cotolog No, 



5. Report Dale 

April. 1974 



6. PerforminB Orgoniialion Code 



8. Performing Orgoniiation Report No. 



10. Work Unit No. 



9, Perlorming Orgonliotion Nome and Addro»» 
SCITRAN 
liox 5456 

Santa pfrrhara, CA 931 Q S 

12. Sponsoring Aaoney Name and Address 



'^a"io^aril^ronautics"ar,d Space Adialnlatration 
Washington, D.C. 20546 



11. Contract or Gfont No. 

NASw-2483 



13. Type of Report end Period Covered 

Translation 



14. Sponsoring Ajooey Code 



15. Supplementary Notei ,, -. 

Translation of "Die Landelasten der Do 3l bei Vertikallandungen und 
Eolgerungqn filr zukiinftige VSTOL Entwicklungen" , Dornier-Werke G.m.b.H. , 
Friedrichsha^ten (W. Ger.), BMVg-EBWT-72-24; 1972, 72 pp. 



16. Abitreet 

The results of 83 vertical landings carried out during^ the Do 31 VSTOL 
Experimental Program are reported. During 23 landings, undercarriage 
reactions, as well as sinking speeds were measured; of the remaining 60 
landings only sinking speeds could be evaluate^. Undercarriage reaction 
factors and sinking speeds are plotted as frequency .distributions and 
discussed. The result of the evaluation of the landing experiments can 
be summarized as follows: VTOL airplanes of a conception similar to. the 
Do 31 and using manual control during the end of descent would experience 
considerably higher sinking speeds than conventional aircraft. Further 
more, a typical jumping of the airplane after touchdown and a following 
second Impact prove unfavorable, this second impact resulting in higher 
undercarriage reactions than the first one. The horizontal loads occurring 
with vertical landings are smaller than expected. 



17. Key Word* (Selected by Author{i)) 



IB. Distribution Slolement 



Unclassified - Unlimited 



19. Seeurtty Clossif. (o* ttili report) 

Unclassified 



20. Security Clofslf. (of this page) 

Unclassified 



21. No. of Paget 

52 



22. Price 



53